CS 27 Amdt 2 Final

Annex to ED Decision 2008/009/R European Aviation Safety Agency

Certification Specifications

for

Small Rotorcraft

CS 27

Amendment 2 17 November 2008 Amendment 2

CS 27 Annex to ED Decision 2008/009/R CONTENTS (general layout) CS–27 SMALL ROTORCRAFT BOOK 1 – AIRWORTHINESS CODE SUBPART A – GENERAL SUBPART B – FLIGHT SUBPART C – STRENGTH REQUIREMENTS SUBPART D – DESIGN AND CONSTRUCTION SUBPART E – POWERPLANT SUBPART F – EQUIPMENT SUBPART G – OPERATING LIMITATIONS AND INFORMATION APPENDICES: A, B and C BOOK 2 – ACCEPTABLE MEANS OF COMPLIANCE (AMC): AMCs Amendment 2 C 1

Annex to ED Decision 2008/009/R CS2 7 PREAMBLE CS2 7 Amendment 1 Effective: 30/11/2007 The following is a list of paragraphs affected by this amendment.

Book 1 Subpart B ∑ CS 27.25 Amended (NPA 11/2006) ∑ CS 27.49 Created by renaming CS 27.73 (NPA 11/2006) ∑ CS 27.51 Amended (NPA 11/2006) ∑ CS 27.73 Deleted and moved to CS 27.49 (NPA 11/2006) ∑ CS 27.75 Amended (NPA 11/2006) ∑ CS 27.79 Amended (NPA 11/2006) ∑ CS 27.143 Amended (NPA 11/2006) ∑ CS 27.173 Amended (NPA 11/2006) ∑ CS 27.175 Amended (NPA 11/2006) ∑ CS 27.177 Amended (NPA 11/2006) Subpart E ∑ CS 27.903 Amended (NPA 11/2006) Subpart G ∑ CS 27.1587 Amended (NPA 11/2006) Appendices ∑ CS2 7 Appendix B Amended (NPA 11/2006) CS2 7 Amendment 2 Effective: 17/11/2008 The following is a list of paragraphs affected by this amendment.

Book 1 Subpart F ∑ CS 27.1305 Amended (NPA 20071 7) Appendices ∑ CS2 7 Appendix A Amended (NPA 20071 7) ∑ CS2 7 Appendix C Amended (NPA 20071 7) Book 2 ∑ AMC 27 General Amended (NPA 20071 7) ∑ AMC 27.351 Created (NPA 20071 7) ∑ AMC 27.602 Deleted (NPA 20071 7) ∑ AMC 27.865 Created (NPA 20071 7) ∑ AMC 27.1305(t) and (u) Deleted (NPA 20071 7) ∑ AMC MG4 Created (NPA 20071 7) Amendment 2 P1

Annex to ED Decision 2008/009/R CS 27 BOOK 1

EASA Certification Specifications

for

SMALL ROTORCRAFT

CS 27

Book 1

Airworthiness code

Amendment 2 1 0 1 Annex to ED Decision 2008/009/R CS–27 BOOK 1 SUBPART A – GENERAL CS 27.1 Applicability (a) This Airworthiness Code is applicable to small rotorcraft with maximum weights of 3 175 kg (7 000 lbs) or less and nine or less passenger seats.

(b) reserved (c) Multi engine rotorcraft may be type certificated as Category A provided the requirements referenced in Appendix C are met.

JAR 27.2 Special Retroactive Requirements (a) reserved (b) For rotorcraft with a certification basis established prior to 1 May 2001 (1) The maximum passenger seat capacity may be increased to eight or nine provided compliance is shown with all the INTENTIONALLY LEFT BLANK airworthiness requirements of this initial issue of CS27.

(2) The maximum weight may be increased to greater than 2 722 kg (6 000 lbs) provided (i) The number of passenger seats is not increased above the maximum number previously certificated; or (ii) Compliance is shown with all of the airworthiness requirements of this initial issue of CS 27.

Amendment 2 1A1 Annex to ED Decision 2008/009/R CS–27 BOOK 1 SUBPART B – FLIGHT GENERAL (iii) The weight of full oil capacity; and (iv) For each seat, an occupant CS 27.21 Proof of compliance weight of 77 kg (170 lbs) or any lower weight for which certification is Each requirement of this Subpart must be met requested.

at each appropriate combination of weight and centre of gravity within the range of loading (b) Minimum weight. The minimum weight, conditions for which certification is requested.

the lowest weight at which compliance with each This must be shown: applicable requirement of this CS–27 is shown, must be established so that it is: (a) By tests upon a rotorcraft of the type for which certification is requested or by calculations (1) Not more than the sum of: based on, and equal in accuracy to, the results of (i) The empty weight determined testing; and under CS 27.29; and (b) By systematic investigation of each (ii) The weight of the minimum required combination of weight and centre of crew necessary to operate the rotorcraft, gravity if compliance cannot be reasonably assuming for each crew member a weight inferred from combinations investigated.

no more than 77 kg (170 lbs), or any lower weight selected by the applicant or included in the loading instructions; and CS 27.25 Weight limits (2) Not less than: (a) Maximum weight. The maximum (i) The lowest weight selected by weight, the highest weight at which compliance the applicant; with each applicable requirement of this CS–27 is shown, must be established so that it is: (ii) The design minimum weight, the lowest weight at which compliance (1) Not more than: with each applicable structural loading (i) The highest weight selected condition of this CS–27 is shown; or by the applicant; (iii) The lowest weight at which (ii) The design maximum weight, compliance with each applicable flight the highest weight at which compliance requirement of this CS–27 is shown.

with each applicable structural loading (c) Total weight with jettisonable external condition of this CS–27 is shown; load . A total weight for the rotorcraft with a jettisonable external load attached that is greater (iii) The highest weight at which than the maximum weight established under sub compliance with each applicable flight paragraph (a) may be established for any requirement of this CS–27 is shown; or rotorcraft load combination if: (iv) The highest weight, as a (1) The rotorcraft load combination function of altitude and temperature, in does not include human external cargo, which the provisions of CS 27.79 and/or CS 27.143(c)(1) are demonstrated if the (2) Structural component approval for operating conditions (altitude and external load operations under either CS temperature) prescribed by those 27.865, or under equivalent operational requirements can not be met; and standards is obtained, (2) Not less than the sum of: (3) The portion of the total weight that is greater than the maximum weight (i) The empty weight determined established under sub paragraph (a) is made under CS 27.29; up only of the weight of all or part of the (ii) The weight of usable fuel jettisonable external load, appropriate to the intended operation (4) Structural components of the with full payload; rotorcraft are shown to comply with the applicable structural requirements of this CS Amendment 2 1–B– 1 Annex to ED Decision 2008/009/R CS–27 BOOK 1 27 under the increased loads and stresses caused by the weight increase over that CS 27.33 Main rotor speed and pitch established under sub paragraph (a) , and limits (5) Operation of the rotorcraft at a (a) Main rotor speed limits . A range of total weight greater than the maximum main rotor speeds must be established that: certificated weight established under sub (1) With power on, provides adequate paragraph (a) is limited by appropriate margin to accommodate the variations in rotor operating limitations under CS 27.865 (a) and speed occurring in any appropriate manoeuvre, (d).

and is consistent with the kind of governor or [Amdt. No.: 27/1] synchroniser used; and (2) With power off, allows each CS 27.27 Centre of gravity limits appropriate autorotative manoeuvre to be performed throughout the ranges of airspeed The extreme forward and aft centres of gravity and weight for which certification is and, where critical, the extreme lateral centres of requested.

gravity must be established for each weight (b) Normal main rotor high pitch limits established under CS 27.25. Such an extreme (power on). For rotorcraft, except helicopters may not lie beyond: required to have a main rotor low speed warning (a) The extremes selected by the applicant; under sub paragraph (e). It must be shown, with power on and without exceeding approved engine (b) The extremes within which the structure maximum limitations, that main rotor speeds is proven; or substantially less than the minimum approved (c) The extremes within which compliance main rotor speed will not occur under any with the applicable flight requirements is shown.

sustained flight condition. This must be met by: (1) Appropriate setting of the main rotor high pitch stop; CS 27.29 Empty weight and corresponding centre of gravity (2) Inherent rotorcraft characteristics that make unsafe low main rotor speeds (a) The empty weight and corresponding unlikely; or centre of gravity must be determined by weighing the rotorcraft without the crew and payload but (3) Adequate means to warn the pilot with: of unsafe main rotor speeds.

(1) Fixed ballast; (c) Normal main rotor low pitch limits (power off). It must be shown, with power off, (2) Unusable fuel; and that: (3) Full operating fluids, including: (1) The normal main rotor low pitch (i) Oil; limit provides sufficient rotor speed, in any autorotative condition, under the most critical (ii) Hydraulic fluid; and combinations of weight and airspeed; and (iii) Other fluids required for (2) It is possible to prevent normal operation of rotorcraft systems, overspeeding of the rotor without exceptional except water intended for injection in the piloting skill.

engines.

(d) Emergency high pitch . If the main rotor (b) The condition of the rotorcraft at the high pitch stop is set to meet sub paragraph time of determining empty weight must be one (b)(1), and if that stop cannot be exceeded that is well defined and can be easily repeated, inadvertently, additional pitch may be made particularly with respect to the weights of fuel, available for emergency use.

oil, coolant, and installed equipment.

(e) Main rotor low speed warning for helicopters . For each single engine helicopter, and each multie ngine helicopter that does not have an CS 27.31 Removable ballast approved device that automatically increases power Removable ballast may be used in showing on the operating engines when one engine fails, compliance with the flight requirements of this there must be a main rotor low speed warning which Subpart.

meets the following requirements: Amendment 2 1–B– 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (1) The warning must be furnished to (1) 80%, at and below standard the pilot in all flight conditions, including temperature; and power on and power off flight, when the speed (2) 34%, at and above standard of a main rotor approaches a value that can temperature plus 28°C (50°F) between these jeopardise safe flight.

two temperatures, the relative humidity must (2) The warning may be furnished vary linearly.

either through the inherent aerodynamic (f) For turbine enginepo wered rotorcraft, a qualities of the helicopter or by a device.

means must be provided to permit the pilot to (3) The warning must be clear and determine prior to takeo ff that each engine is distinct under all conditions, and must be capable of developing the power necessary to clearly distinguishable from all other achieve the applicable rotorcraft performance warnings. A visual device that requires the prescribed in this Subpart.

attention of the crew within the cockpit is not acceptable by itself.

CS 27.49 Performance at minimum (4) If a warning device is used, the operating speed device must automatically de activate and reset when the low speed condition is (a) For helicopters: corrected. If the device has an audible (1) The hovering ceiling must be warning, it must also be equipped with a determined over the ranges of weight, altitude, means for the pilot to manually silence the and temperature for which certification is audible warning before the low speed requested, with: condition is corrected.

(i) Take off power; (ii) The landing gear extended; PERFORMANCE and (iii) The helicopter in ground effect at a height consistent with normal CS 27.45 General take off procedures; and (a) Unless otherwise prescribed, the (2) The hovering ceiling determined in performance requirements of this Subpart must sub paragraph (a)(1) of this paragraph must be be met for still air and a standard atmosphere.

at least: (b) The performance must correspond to the (i) For reciprocating engine engine power available under the particular powered helicopters, 1219 m (4 000 ft) at ambient atmospheric conditions, the particular maximum weight with a standard flight condition, and the relative humidity atmosphere; or specified in sub paragraphs (d) or (e), as appropriate. (ii) For turbine engine powered helicopters, 762 m (2 500 ft) pressure (c) The available power must correspond to altitude at maximum weight at a engine power, not exceeding the approved power, temperature of standard +22°C (+40°F).

less: (3) The out of ground effect hovering (1) Installation losses; and performance must be determined over the (2) The power absorbed by the ranges of weight, altitude, and temperature for accessories and services appropriate to the which certification is requested, using take off particular ambient atmospheric conditions and power.

the particular flight condition.

(b) For rotorcraft other than helicopters, the (d) For reciprocating engine powered steady rate of climb at the minimum operating rotorcraft, the performance, as affected by engine speed must be determined, over the ranges of power, must be based on a relative humidity of weight, altitude, and temperature for which 80% in a standard atmosphere. certification is requested, with: (e) For turbine engine powered rotorcraft, (1) Take off power; and the performance, as affected by engine power, (2) The landing gear extended.

must be based on a relative humidity of: Amendment 2 1–B– 3 Annex to ED Decision 2008/009/R CS–27 BOOK 1 [Amdt. No.: 27/1] (ii) At maximum weight; and (iii) With maximum continuous power on each engine.

CS 27.51 Take off (2) The steady rate of climb must be The takeo ff, with takeo ff power and rpm at the determined: most critical center of gravity, and with weight (i) At the climb speed selected by from the maximum weight at sea level to the the applicant at or below V ; weight for which take off certification is NE requested for each altitude covered by this (ii) Within the range from sea paragraph: level up to the maximum altitude for which certification is requested; (a) May not require exceptional piloting skill or exceptionally favourable conditionsthroughout (iii) For the weights and the ranges of altitude from standard seal evel temperatures that correspond to the conditions to the maximum altitude for which take altitude range set forth in sub paragraph off and landing certification is requested, and (b)(2)(ii) and for which certification is requested; and (b) Must be made in such a manner that a landing can be made safely at any point along the (iv) With maximum continuous flight path if an engine fails. This must be power on each engine.

demonstrated up to the maximum altitude for which takeo ff and landing certification is requested or 2134m (7,000 ft) density altitude, whichever is less.

CS 27.67 Climb: one engine inoperative [Amdt. No.: 27/1] For multi engine helicopters, the steady rate of climb (or descent), at V (or at the speed for Y minimum rate of descent), must be determined CS 27.65 Climb: all engines operating with: (a) For rotorcraft other than helicopters: (a) Maximum weight; (1) The steady rate of climb, at V Y (b) The critical engine inoperative and the must be determined: remaining engines at either: (i) With maximum continuous (1) Maximum continuous power and, power on each engine; for helicopters for which certification for the use of 30 minute one engine inoperative (OEI) (ii) With the landing gear power is requested, at 30 minute OEI power; retracted; and or (iii) For the weights, altitudes, and (2) Continuous OEI power for temperatures for which certification is helicopters for which certification for the use requested; and of continuous OEI power is requested.

(2) The climb gradient, at the rate of climb determined in accordance with sub CS 27.71 Glide performance paragraph (a)(1), must be either: For single engine helicopters and multi engine (i) At least 1:10 if the horizontal helicopters that do not meet the category A distance required to take off and climb engine isolation requirements of CS–27, the over a 15 m (50 ft) obstacle is minimum rate of descent airspeed and the best determined for each weight, altitude, and angle of glide airspeed must be determined in temperature within the range for which autorotation at: certification is requested; or (a) Maximum weight; and (ii) At least 1:6 under standard sea level conditions.

(b) Rotor speed(s) selected by the applicant.

(b) Each helicopter must meet the following requirements: CS 27.75 Landing (1) V must be determined: Y (a) The rotorcraft must be able to be landed (i) For standard sea level with no excessive vertical acceleration, no conditions; Amendment 2 1–B– 4 Annex to ED Decision 2008/009/R CS–27 BOOK 1 tendency to bounce, nose over, ground loop, limits and at the minimum installed porpoise, or water loop, and without exceptional specification power available for the most piloting skill or exceptionally favourable critical combination of approved ambient conditions, with: temperature and pressure altitude resulting in 2134m (7000 ft) density altitude or the (1) Approach or autorotation speeds maximum altitude capability of the helicopter, appropriate to the type of rotorcraft and whichever is less; and selected by the applicant; (3) For other rotorcraft, conditions (2) The approach and landing made appropriate to the type.

with: (i) Power off, for single engine [Amdt. No.: 27/1] rotorcraft and entered from steady state autorotation; or FLIGHT CHARACTERISTICS (ii) One engine inoperative (OEI) for multi engine rotorcraft with each operating engine within approved operating limitations, and entered from CS 27.141 General an established OEI approach.; The rotorcraft must: (b) Multi engine rotorcraft must be able to (a) Except as specifically required in the be landed safely after complete power failure applicable paragraph, meet the flight under normal operating conditions.

characteristics requirements of this Subpart: [Amdt. No.: 27/1] (1) At the altitudes and temperatures expected in operation; CS 27.79 Limiting height speed envelope (2) Under any critical loading condition within the range of weights and (a) If there is any combination of height and centres of gravity for which certification is forward speed, including hover, under which a requested; safe landing cannot be made under the applicable power failure condition in sub paragraph (b), a (3) For power on operations, under any limiting height speed envelope must be condition of speed, power, and rotor rpm for established, including all pertinent information, which certification is requested; and for that condition, throughout the ranges of: (4) For power off operations, under (1) Altitude, from standard sea level any condition of speed and rotor rpm for conditions to the maximum altitude capability which certification is requested that is of the rotorcraft, or 2134 m (7 000 ft) density attainable with the controls rigged in altitude, whichever is less; and accordance with the approved rigging instructions and tolerances; (2) Weight from the maximum weight at sea level to the weight selected by the (b) Be able to maintain any required flight applicant for each altitude covered by sub condition and make a smooth transition from any paragraph (a)(1) of this paragraph. For flight condition to any other flight condition helicopters, the weight at altitudes above sea without exceptional piloting skill, alertness, or level may not be less than the maximum strength, and without danger of exceeding the weight or the highest weight allowing limit load factor under any operating condition hovering out of ground effect whichever is probable for the type, including: lower.

(1) Sudden failure of one engine, for (b) The applicable power failure conditions multi engine rotorcraft meeting category A are: engine isolation requirements of CS–29; (1) For single engine helicopters, full (2) Sudden, complete power failure for autorotation; other rotorcraft; and (2) For multi engine helicopters, OEI, (3) Sudden, complete control system where engine isolation features ensure failures specified in CS 27.695; and continued operation of the remaining engines, (c) Have any additional characteristic and the remaining engine(s) within approved required for night or instrument operation, if Amendment 2 1–B– 5 Annex to ED Decision 2008/009/R CS–27 BOOK 1 certification for those kinds of operation is (i) Weight selected by the requested. Requirements for helicopter applicant; instrument flight are contained in appendix B.

(ii) Critical centre of gravity; and (iii) Critical rotor rpm.

CS 27.143 Controllability and (d) Wind velocities from zero to at least 31 manoeuvrability km/h (17 knots), from all azimuths, must be (a) The rotorcraft must be safely controllable established in which the rotorcraft can be and manoeuvrable: operated without loss of control out of ground effect, with: (1) During steady flight; and (1) Weight selected by the applicant; (2) During any manoeuvre appropriate to the type, including: (2) Critical centre of gravity; (i) Take off; (3) Rotor rpm selected by the applicant; and (ii) Climb; (4) Altitude, from standard sea level (iii) Level flight; conditions to the maximum take off and (iv) Turning flight; landing altitude capability of the rotorcraft.

(v) Autorotation; (e) The rotorcraft, after (vi) Landing (power on and (1) failure of one engine in the case of power off); and multi engine rotorcraft that meet Category A engine isolation requirements, or (vii) Recovery to power on flight from a balked autorotative approach.

(2) complete engine failure in the case of other rotorcraft, must be controllable over (b) The margin of cyclic control must allow the range of speeds and altitudes for which satisfactory roll and pitch control at V with: NE certification is requested when such power (1) Critical weight; failure occurs with maximum continuous power and critical weight. No corrective (2) Critical centre of gravity; action time delay for any condition following (3) Critical rotor rpm; and power failure may be less than: (4) Power off, except for helicopters (i) For the cruise condition, one demonstrating compliance with sub paragraph second, or normal pilot reaction time (f), and power on.

(whichever is greater); and (c) Wind velocities from zero to at least 31 (ii) For any other condition, km/h (17 knots), from all azimuths, must be normal pilot reaction time.

established in which the rotorcraft can be (f) For helicopters for which a V (power NE operated without loss of control on or near the off) is established under CS 27.1505 (c), ground in any manoeuvre appropriate to the type, compliance must be demonstrated with the such as crosswind take offs, sideward flight and following requirements with critical weight, rearward flight: critical centre of gravity, and critical rotor rpm: (1) With altitude, from standard sea (1) The helicopter must be safely level conditions to the maximum take off and slowed to V NE (power off), without landing altitude capability of the rotorcraft or exceptional pilot skill, after the last operating 2134m (7000 ft) density altitude, whichever is engine is made inoperative at power on V NE ; less; with: (2) At a speed of 1.1 V (power off), NE (i) Critical weight; the margin of cyclic control must allow (ii) Critical centre of gravity; and satisfactory roll and pitch control with power off.

(iii) Critical rotor rpm.

[Amdt. No.: 27/1] (2) For take off and landing altitudes above 2134m (7000 ft) density altitude with: CS 27.151 Flight controls Amendment 2 1–B– 6 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (a) Longitudinal, lateral, directional, and (a) Climb . Static longitudinal stability must collective controls may not exhibit excessive be shown in the climb condition at speeds breakout force, friction or preload. fromVy 19 km/h (10 knots) to Vy + 19 km/h (10 knots), with: (b) Control system forces and free play may not inhibit a smooth, direct rotorcraft response to (1) Critical weight; control system input.

(2) Critical centre of gravity; (3) Maximum continuous power; CS 27.161 Trim control (4) The landing gear retracted; and The trim control: (5) The rotorcraft trimmed at V .

Y (a) Must trim any steady longitudinal, (b) Cruise . Static longitudinal stability must lateral, and collective control forces to zero in be shown in the cruise condition at speeds from level flight at any appropriate speed; and 0.8 V 19 km/h (10 knots) to 0.8 V + 19 NE NE (b) May not introduce any undesirable km/h (10 knots) or, if V is less than 0.8 V , H NE discontinuities in control force gradients. from V 19 km/h (10 knots) to V + 19 km/h H H (10 knots), with: (1) Critical weight; CS 27.171 Stability: general (2) Critical centre of gravity; The rotorcraft must be able to be flown, without undue pilot fatigue or strain, in any (3) Power for level flight at 0.8 V or NE normal manoeuvre for a period of time as long as V , whichever is less; H that expected in normal operation. At least three (4) The landing gear retracted; and landings and take offs must be made during this demonstration. (5) The rotorcraft trimmed at0.8 V NE or V , whichever is less.

H (c) V . Static longitudinal stability must be NE CS 27.173 Static longitudinal stability shown at speeds from V – 28 km/h (20 knots) NE (a) The longitudinal control must be to V with: NE designed so that a rearward movement of the (1) Critical weight; control is necessary to obtain an airspeed less than the trim speed, and a forward movement of (2) Critical centre of gravity; the control is necessary to obtain an airspeed (3) Power required for level flight at more than the trim speed.

V – 19 km/h (10 knots) or maximum NE (b) Throughout the full range of altitude for continuous power, whichever is less; which certification is requested,with the throttle (4) The landing gear retracted; and and collective pitch held constant during the manoeuvres specified in CS 27.175(a) through (5) The rotorcraft trimmed at V – 19 NE (d), the slope of the control position versus km/h (10 knots).

airspeed curve must be positive. However, in (d) Autorotation . Static longitudinal limited flight conditions or modes of operation stability must be shown in autorotation at: determined by the Agency to be acceptable, the slope of the control position versus airspeed (1) Airspeeds from the minimum rate curve may be neutral or negative if the rotorcraft of descent airspeed – 19 km/h (10 knots) to the possesses flight characteristics that allow the minimum rate of descent airspeed + 19 km/h pilot to maintain airspeed within ±9 km/h (±5 (10 knots), with: knots) of the desired trim airspeed without (i) Critical weight; exceptional piloting skill or alertness.

(ii) Critical centre of gravity; [Amdt. No.: 27/1] (iii) The landing gear extended; and CS 27.175 Demonstration of static (iv) The rotorcraft trimmed at the longitudinal stability minimum rate of descent airspeed.

Amendment 2 1–B– 7 Annex to ED Decision 2008/009/R CS–27 BOOK 1 GROUND AND WATER HANDLING (2) Airspeeds from the best angle of CHARACTERISTICS glide airspeed – 19 km/h (10 knots) to the best angle of glide airspeed + 19 km/h (10 knots), with: CS 27.231 General (i) Critical weight; The rotorcraft must have satisfactory ground (ii) Critical centre of gravity; and water handling characteristics, including (iii) The landing gear retracted; freedom from uncontrollable tendencies in any and condition expected in operation.

(iv) The rotorcraft trimmed at the best angle of glide airspeed.

CS 27.235 Taxying condition [Amdt. No.: 27/1] The rotorcraft must be designed to withstand the loads that would occur when the rotorcraft is taxied over the roughest ground that may CS 27.177 Static directional stability reasonably be expected in normal operation.

(a) The directional controls must operate in such a manner that the sense and direction of motion of the rotorcraft following control CS 27.239 Spray characteristics displacement are in the direction of the pedal If certification for water operation is motion with throttle and collective controls held requested, no spray characteristics during constant at the trim conditions specified in CS taxying, take off, or landing may obscure the 27.175 (a), (b), and (c). Sideslip angles must vision of the pilot or damage the rotors, increase with steadily increasing directional propellers, or other parts of the rotorcraft.

control deflection for sideslip angles up to the lesser of: (1) ± 25 degrees from trim at a speed of CS 27.241 Ground resonance 28 km/h (15 knots) less than the speed for The rotorcraft may have no dangerous minimum rate of descent varying linearly to tendency to oscillate on the ground with the rotor ± 10 degrees from trim at V ; NE turning.

(2) The steady state sideslip angles established by CS 27.351; MISCELLANEOUS FLIGHT REQUIREMENTS (3) A sideslip angle selected by the applicant which corresponds to a sideforce of at least 0.1g; or, CS 27.251 Vibration (4) The sideslip angle attained by maximum directional control input. Each part of the rotorcraft must be free from excessive vibration under each appropriate speed (b) Sufficient cues must accompany the and power condition.

sideslip to alert the pilot when approaching sideslip limits.

(c) During the manoeuvre specified in sub paragraph (a) of this paragraph, the sideslip angle versus directional control position curve may have a negative slope within a small range of angles around trim, provided the desired heading can be maintained without exceptional piloting skill or alertness.

[Amdt. No.: 27/1] Amendment 2 1–B– 8 Annex to ED Decision 2008/009/R CS–27 BOOK 1 SUBPART C – STRENGTH REQUIREMENTS GENERAL analysis (static or fatigue) may be used only if the structure conforms to those structures for which experience has shown this method to be reliable.

CS 27.301 Loads In other cases, substantiating load tests must be (a) Strength requirements are specified in made.

terms of limit loads (the maximum loads to be (b) Proof of compliance with the strength expected in service) and ultimate loads (limit requirements of this Subpart must include: loads multiplied by prescribed factors of safety).

Unless otherwise provided, prescribed loads are (1) Dynamic and endurance tests of limit loads.

rotors, rotor drives, and rotor controls; (b) Unless otherwise provided, the specified (2) Limit load tests of the control air, ground, and water loads must be placed in system, including control surfaces; equilibrium with inertia forces, considering each (3) Operation tests of the control item of mass in the rotorcraft. These loads must system; be distributed to closely approximate or conservatively represent actual conditions.

(4) Flight stress measurement tests; (c) If deflections under load would (5) Landing gear drop tests; and significantly change the distribution of external (6) Any additional tests required for or internal loads, this redistribution must be new or unusual design features.

taken into account.

CS 27.303 Factor of safety CS 27.309 Design limitations Unless otherwise provided, a factor of safety of The following values and limitations must be 1.5 must be used. This factor applies to external established to show compliance with the and inertia loads unless its application to the structural requirements of this Subpart: resulting internal stresses is more conservative.

(a) The design maximum weight.

(b) The main rotor rpm ranges power on and CS 27.305 Strength and deformation power off.

(a) The structure must be able to support (c) The maximum forward speeds for each limit loads without detrimental or permanent main rotor rpm within the ranges determined in deformation. At any load up to limit loads, the sub paragraph (b) .

deformation may not interfere with safe operation.

(d) The maximum rearward and sideward flight speeds.

(b) The structure must be able to support ultimate loads without failure. This must be (e) The centre of gravity limits shown by: corresponding to the limitations determined under sub paragraphs (b), (c), and (d) .

(1) Applying ultimate loads to the structure in a static test for at least 3 seconds; (f) The rotational speed ratios between each or powerplant and each connected rotating component.

(2) Dynamic tests simulating actual load application.

(g) The positive and negative limit manoeuvring load factors.

CS 27.307 Proof of structure (a) Compliance with the strength and FLIGHT LOADS deformation requirements of this Subpart must be CS 27.321 General shown for each critical loading condition accounting for the environment to which the (a) The flight load factor must be assumed structure will be exposed in operation. Structural to act normal to the longitudinal axis of the Amendment 2 1–C–1 Annex to ED Decision 2008/009/R CS–27 BOOK 1 rotorcraft, and to be equal in magnitude and a = The angle between the projection, in the opposite in direction to the rotorcraft inertia load plane of symmetry, of the axis of no factor at the centre of gravity.

feathering and a line perpendicular to the (b) Compliance with the flight load flight path (positive when the axis is requirements of this Subpart must be shown: pointing aft); (1) At each weight from the design Ω = The angular velocity of rotor; and minimum weight to the design maximum R = The rotor radius.

weight; and (2) With any practical distribution of disposable load within the operating CS 27.341 Gust loads limitations in the Rotorcraft Flight Manual.

The rotorcraft must be designed to withstand, at each critical airspeed including hovering, the loads resulting from a vertical gust of 9.1 m/s (30 CS 27.337 Limit manoeuvring load factor ft/s).

The rotorcraft must be designed for: (a) A limit manoeuvring load factor ranging CS 27.351 Yawing conditions from a positive limit of 3.5 to a negative limit of – 1.0; or (a) Each rotorcaft must be designed for the loads resulting from the manoeuvres specified in (b) Any positive limit manoeuvring load factor sub paragraphs (b) and (c) with: not less than 2.0 and any negative limit manoeuvring load factor of not less than – 0.5 for (1) Unbalanced aerodynamic moments which: about the centre of gravity which the aircraft reacts to in a rational or conservative manner (1) The probability of being exceeded considering the principal masses furnishing is shown by analysis and flight tests to be the reacting inertia forces; and extremely remote; and (2) Maximum main rotor speed.

(2) The selected values are appropriate to each weight condition between the design (b) To produce the load required in sub maximum and design minimum weights.

paragraph (a) , in unaccelerated flight with zero yaw, at forward speeds from zero up to 0.6 V : NE (1) Displace the cockpit directional CS 27.339 Resultant limit manoeuvring control suddenly to the maximum deflection loads limited by the control stops or by the maximum pilot force specified in CS 27.397 The loads resulting from the application of (a); limit manoeuvring load factors are assumed to act at the centre of each rotor hub and at each (2) Attain a resulting sideslip angle or auxiliary lifting surface, and to act in directions, 90°, whichever is less; and and with distributions of load among the rotors (3) Return the directional control and auxiliary lifting surfaces, so as to represent suddenly to neutral.

each critical manoeuvring condition, including power on and power off flight with the maximum (c) To produce the load required in sub design rotor tip speed ratio. The rotor tip speed paragraph (a) , in unaccelerated flight with zero ratio is the ratio of the rotorcraft flight velocity yaw, at forward speeds from 0.6 V up to V or NE NE component in the plane of the rotor disc to the V , whichever is less: H rotational tip speed of the rotor blades, and is (1) Displace the cockpit directional expressed as follows: control suddenly to the maximum deflection cosa V limited by the control stops or by the μ = ΩR maximum pilot force specified in CS 27.397 where: (a); V = The airspeed along the flight path; Amendment 2 1–C–2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (2) Attain a resulting sideslip angle or allows the pilot to apply, but not less than 0.60 15°, whichever is less, at the lesser speed of times the forces specified in CS 27.397.

V or V ; NE H (b) Each primary control system including (3) Vary the sideslip angles of sub its supporting structure, must be designed as paragraphs (b)(2) and (c)(2) directly with follows: speed; and (1) The system must withstand loads (4) Return the directional control resulting from the limit pilot forces prescribed suddenly to neutral. in CS 27.397.

(2) Notwithstanding subpa ragraph (b)(3) , when power operated actuator controls CS 27.361 Engine torque or power boost controls are used, the system (a) For turbine engines, the limit torque may must also withstand the loads resulting from not be less than the highest of: the force output of each normally energised power device, including any single power (1) The mean torque for maximum boost or actuator system failure.

continuous power multiplied by 1.25; (3) If the system design or the normal (2) The torque required by CS 27.923; operating loads are such that a part of the (3) The torque required by CS 27.927; system cannot react to the limit forces or prescribed in CS 27.397, that part of the system must be designed to withstand the (4) The torque imposed by sudden maximum loads that can be obtained in engine stoppage due to malfunction or normal operation. The minimum design loads structural failure (such as compressor must, in any case, provide a rugged system for jamming).

service use, including consideration of fatigue, (b) For reciprocating engines, the limit jamming, ground gusts, control inertia and torque may not be less than the mean torque for friction loads. In the absence of rational maximum continuous power multiplied by: analysis, the design loads resulting from 0.60 of the specified limit pilot forces are (1) 1.33, for engines with five or more acceptable minimum design loads.

cylinders; and (4) If operational loads may be (2) Two, three, and four, for engines exceeded through jamming, ground gusts, with four, three, and two cylinders, control inertia, or friction, the system must respectively.

withstand the limit pilot forces specified in CS 27.397, without yielding.

CONTROL SURFACE AND SYSTEM LOADS CS 27.397 Limit pilot forces and torques CS 27.391 General (a) Except as provided in sub paragraph (b) Each auxiliary rotor, each fixed or movable the limit pilot forces are as follows: stabilising or control surface, and each system (1) For foot controls, 578 N (130 lbs).

operating any flight control must meet the requirements of CS 27.395, 27.397, 27.399, (2) For stick controls, 445 N (100 lbs) 27.411 and 27.427.

fore and aft, and 298 N (67 lbs) laterally.

(b) For flap, tab, stabiliser, rotor brake, and CS 27.395 Control system landing gear operating controls, the following (a) The part of each control system from the apply: pilot’s controls to the control stops must be designed (1) Crank, wheel, and lever controls, to withstand pilot forces of not less than – (25.4 + R) x 2.919 N, where R = radius in ˘ È + R 1 (1) The forces specified in CS 27.397; millimetres ( x 50 lbs, where R = ˙ Í ˚ Î or radius in inches), but not less than 222 N (50 (2) If the system prevents the pilot lbs) nor more than 445 N (100 lbs) for hand from applying the limit pilot forces to the operated controls or 578 N (130 lbs) for foot system, the maximum forces that the system Amendment 2 1–C–3 Annex to ED Decision 2008/009/R CS–27 BOOK 1 operated controls, applied at any angle within supporting structure must be designed for the 20° of the plane of motion of the control. combined vertical and horizontal surface loads resulting from each prescribed flight condition, (2) Twist controls, 356 x R Newton considered separately. The flight conditions millimetres, where R = radius in millimetres (80 must be selected so the maximum design loads x R inchpo unds where R = radius in inches).

are obtained on each surface. In the absence of more rational data, the unsymmetrical horizontal tail surface loading distributions described in this CS 27.399 Dual control system paragraph must be assumed.

Each dual primary flight control system must be designed to withstand the loads that result when pilot forces of 0.75 times those obtained GROUND LOADS under CS 27.395 are applied – CS 27.471 General (a) In opposition; and (a) Loads and equilibrium. For limit ground (b) In the same direction.

loads – (1) The limit ground loads obtained in the landing conditions in this Subpart must be CS 27.411 Ground clearance: tail rotor guard considered to be external loads that would occur in the rotorcraft structure if it were (a) It must be impossible for the tail rotor to acting as a rigid body; and contact the landing surface during a normal landing. (2) In each specified landing condition, the external loads must be placed in (b) If a tail rotor guard is required to show equilibrium with linear and angular inertia compliance with sub paragraph (a): loads in a rational or conservative manner.

(1) Suitable design loads must be (b) Critical centres of gravity . The critical established for the guard; and centres of gravity within the range for which (2) The guard and its supporting certification is requested must be selected so that structure must be designed to withstand those the maximum design loads are obtained in each loads. landing gear element.

CS 27.473 Ground loading conditions and CS 27.427 Unsymmetrical loads assumptions (a) Horizontal tail surfaces and their (a) For specified landing conditions, a supporting structure must be designed for design maximum weight must be used that is not unsymmetrical loads arising from yawing and less than the maximum weight. A rotor lift may rotor wake effects in combination with the be assumed to act through the centre of gravity prescribed flight conditions.

throughout the landing impact. This lift may not exceed two thirds of the design maximum (b) To meet the design criteria of sub weight.

paragraph (a), in the absence of more rational data, both of the following must be met: (b) Unless otherwise prescribed, for each specified landing condition, the rotorcraft must (1) 100% of the maximum loading be designed for a limit load factor of not less than from the symmetrical flight conditions acts on the limit inertia load factor substantiated under the surface on one side of the plane of CS 27.725.

symmetry and no loading acts on the other side.

(2) 50% of the maximum loading from CS 27.475 Tyres and shock absorbers the symmetrical flight conditions acts on the Unless otherwise prescribed, for each specified surface on each side of the plane of symmetry landing condition, the tyres must be assumed to but in opposite directions.

be in their static position and the shock absorbers (c) For empennage arrangements where the to be in their most critical position.

horizontal tail surfaces are supported by the vertical tail surfaces, the vertical tail surfaces and Amendment 2 1–C–4 Annex to ED Decision 2008/009/R CS–27 BOOK 1 CS 27.477 Landing gear arrangement and to contact the ground on one aft wheel. In this attitude: Paragraphs CS 27.235, 27.479 to 27.485, and CS 27.493 apply to landing gear with two wheels (a) The vertical load must be the same as aft, and one or more wheels forward, of the that obtained on that side under CS 27.479(b)(l); centre of gravity.

and (b) The unbalanced external loads must be reacted by rotorcraft inertia.

CS 27.479 Level landing conditions (a) Attitudes . Under each of the loading conditions prescribed in sub paragraph (b), the CS 27.485 Lateral drift landing conditions rotorcraft is assumed to be in each of the (a) The rotorcraft is assumed to be in the following level landing attitudes: level landing attitude, with: (1) An attitude in which all wheels (1) Side loads combined with one half contact the ground simultaneously.

of the maximum ground reactions obtained in (2) An attitude in which the aft wheels the level landing conditions of CS 27.479 (b) (1); contact the ground with the forward wheels and just clear of the ground.

(2) The loads obtained under sub (b) Loading conditions . The rotorcraft must paragraph (a)(1) applied: be designed for the following landing loading (i) At the ground contact point; conditions: or (1) Vertical loads applied under (ii) For full swivelling gear, at CS 27.471.

the centre of the axle.

(2) The loads resulting from a (b) The rotorcraft must be designed to combination of the loads applied under sub withstand, at ground contact – paragraph (b)(1) with drag loads at each wheel of not less than 25% of the vertical load at that (1) When only the aft wheels contact wheel.

the ground, side loads of 0.8 times the vertical reaction acting inward on one side, and 0,6 (3) If there are two wheels forward, a times the vertical reaction acting outward on distribution of the loads applied to those the other side, all combined with the vertical wheels under subpa ragraphs (b)(1) and (2) in a loads specified in sub paragraph (a) ; and ratio of 40:60.

(2) When all wheels contact the (c) Pitching moments . Pitching moments ground simultaneously: are assumed to be resisted by: (i) For the aft wheels, the side (1) In the case of the attitude in sub loads specified in sub paragraph (b)(1); paragraph (a)(1), the forward landing gear, and and (ii) For the forward wheels, a side (2) In the case of the attitude in sub load of 0.8 times the vertical reaction paragraph (a)(2), the angular inertia forces.

combined with the vertical load specified in sub paragraph (a).

CS 27.481 Tail down landing conditions (a) The rotorcraft is assumed to be in the CS 27.493 Braked roll conditions maximum nose up attitude allowing ground Under braked roll conditions with the shock clearance by each part of the rotorcraft.

absorbers in their static positions: (b) In this attitude, ground loads are (a) The limit vertical load must be based on assumed to act perpendicular to the ground.

a load factor of at least: (1) 1.33, for the attitude specified in CS 27.483 One wheel landing conditions CS 27.479 (a)(l); and For the one wheel landing condition, the (2) 1.0 for the attitude specified in CS rotorcraft is assumed to be in the level attitude 27.479 (a)(2); and Amendment 2 1–C–5 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (b) The structure must be designed to (f) Side loads in the level landing attitude .

withstand at the ground contact point of each In the attitudes specified in sub paragraphs (b) wheel with brakes, a drag load at least the lesser and (c) the following apply: of: (1) The side loads must be combined at (1) The vertical load multiplied by a each wheel with one half of the maximum coefficient of friction of 0.8; and vertical ground reactions obtained for that wheel under subpa ragraphs (b) and (c). In this (2) The maximum value based on condition the side loads must be: limiting brake torque.

(i) For the forward wheels, 0.8 times the vertical reaction (on one CS 27.497 Ground loading conditions: side) acting inward, and 0.6 times the landing gear with tail wheels vertical reaction (on the other side) acting outward; and (a) General . Rotorcraft with landing gear with two wheels forward, and one wheel aft, of (ii) For the rear wheel, 0.8 times the centre of gravity must be designed for loading the vertical reaction.

conditions as prescribed in this paragraph.

(2) The loads specified in sub paragraph (f)(1) must be applied: (b) Level landing attitude with only the forward wheels contacting the ground. In this (i) At the ground contact point attitude: with the wheel in the trailing position (for non full swivelling landing gear or (1) The vertical loads must be applied for full swivelling landing gear with a under CS 27.471 to 27.475; lock, steering device, or shimmy damper (2) The vertical load at each axle must to keep the wheel in the trailing be combined with a drag load at that axle of position); or not less than 25% of that vertical load; and (ii) At the centre of the axle (for full swivelling landing gear without a (3) Unbalanced pitching moments are lock, steering device, or shimmy assumed to be resisted by angular inertia damper).

forces.

(g) Braked roll conditions in the level (c) Level landing attitude with all wheels landing attitude . In the attitudes specified in contacting the ground simultaneously . In this sub paragraphs (b) and (c), and with shock attitude, the rotorcraft must be designed for absorbers in their static positions, the rotorcraft landing loading conditions as prescribed in sub must be designed for braked roll loads as follows: paragraph (b) .

(1) The limit vertical load must be (d) Maximum nose up attitude with only the based on a limit vertical load factor of not less rear wheel contacting the ground . The attitude than: for this condition must be the maximum nose up attitude expected in normal operation, including (i) 1.0 for the attitude specified autorotative landings. In this attitude: in sub paragraph (b); and (1) The appropriate ground loads (ii) 1.33, for the attitude specified specified in sub paragraphs (b)(1) and (2) in sub paragraph (c).

must be determined and applied, using a (2) For each wheel with brakes, a drag rational method to account for the moment load must be applied, at the ground contact arm between the rear wheel ground reaction point, of not less than the lesser of: and the rotorcraft centre of gravity; or (i) 0.8 times the vertical load; (2) The probability of landing with and initial contact on the rear wheel must be shown to be extremely remote.

(ii) The maximum based on limiting brake torque.

(e) Level landing attitude with only one (h) Rear wheel turning loads in the static forward wheel contacting the ground . In this ground attitude . In the static ground attitude, and attitude, the rotorcraft must be designed for with the shock absorbers and tyres in their static ground loads as specified in sub paragraphs (b) positions, the rotorcraft must be designed for rear (1) and (3).

wheel turning loads as follows: Amendment 2 1–C–6 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (1) A vertical ground reaction equal to rotorcraft contacting the ground along the bottom the static load on the rear wheel must be of both skids, the vertical reactions must be combined with an equal sideload. applied as prescribed in sub paragraph (a).

(2) The load specified in sub (c) Drag reactions in the level landing paragraph (h)(1) must be applied to the rear attitude . In the level attitude, and with the landing gear: rotorcraft contacting the ground along the bottom of both skids, the following apply: (i) Through the axle, if there is a swivel (the rear wheel being assumed to (1) The vertical reactions must be be swivelled 90° to the longitudinal axis combined with horizontal drag reactions of of the rotorcraft); or 50% of the vertical reaction applied at the ground.

(ii) At the ground contact point, if there is a lock, steering device or (2) The resultant ground loads must shimmy damper (the rear wheel being equal the vertical load specified in sub assumed to be in the trailing position).

paragraph (b).

(i) Taxying condition . The rotorcraft and (d) Side loads in the level landing attitude .

its landing gear must be designed for loads that In the level attitude, and with the rotorcraft would occur when the rotorcraft is taxied over contacting the ground along the bottom of both the roughest ground that may reasonably be skids, the following apply: expected in normal operation.

(1) The vertical ground reaction must be: CS 27.501 Ground loading conditions: (i) Equal to the vertical loads landing gear with skids obtained in the condition specified in (a) General . Rotorcraft with landing gear sub paragraph (b); and with skids must be designed for the loading (ii) Divided equally among the conditions specified in this paragraph. In skids.

showing compliance with this paragraph, the following apply: (2) The vertical ground reactions must be combined with a horizontal sideload of 25% (1) The design maximum weight, of their value.

centre of gravity, and load factor must be determined under CS 27.471 to 27.475. (3) The total sideload must be applied equally between the skids and along the length (2) Structural yielding of elastic spring of the skids.

members under limit loads is acceptable.

(4) The unbalanced moments are (3) Design ultimate loads for elastic assumed to be resisted by angular inertia.

spring members need not exceed those obtained in a drop test of the gear with: (5) The skid gear must be investigated for: (i) A drop height of 1.5 times that specified in CS 27.725; and (i) Inward acting sideloads; and (ii) An assumed rotor lift of not (ii) Outward acting sideloads.

more than 1.5 times that used in the (e) One skid landing loads in the level limit drop tests prescribed in CS 27.725.

attitude . In the level attitude, and with the (4) Compliance with subpa ragraphs (b) rotorcraft contacting the ground along the bottom to (e) must be shown with: of one skid only, the following apply: (i) The gear in its most critically (1) The vertical load on the ground deflected position for the landing contact side must be the same as that obtained condition being considered; and on that side in the condition specified in sub paragraph (b).

(ii) The ground reactions rationally distributed along the bottom of (2) The unbalanced moments are the skid tube. assumed to be resisted by angular inertia.

(b) Vertical reactions in the level landing (f) Special conditions . In addition to the attitude . In the level attitude, and with the conditions specified in sub paragraphs (b) and Amendment 2 1–C–7 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (c), the rotorcraft must be designed for the following ground reactions: WATER LOADS (1) A ground reaction load acting up and aft at an angle of 45° to the longitudinal CS 27.521 Float landing conditions axis of the rotorcraft. This load must be: If certification for float operation is requested, (i) Equal to 1.33 times the the rotorcraft, with floats, must be designed to maximum weight; withstand the following loading conditions (where the limit load factor is determined under (ii) Distributed symmetrically CS 27.473 (b) or assumed to be equal to that among the skids; determined for wheel landing gear): (iii) Concentrated at the forward (a) Up load conditions in which: end of the straight part of the skid tube; and (1) A load is applied so that, with the rotorcraft in the static level attitude, the (iv) Applied only to the forward resultant water reaction passes vertically end of the skid tube and its attachment to through the centre of gravity; and the rotorcraft.

(2) The vertical load prescribed in sub (2) With the rotorcraft in the level paragraph (a)(1) is applied simultaneously landing attitude, a vertical ground reaction with an aft component of 0.25 times the load equal to one half of the vertical load vertical component.

determined in sub paragraph (b). This load must be – (b) A side load condition in which: (i) Applied only to the skid tube (1) A vertical load of 0.75 times the and its attachment to the rotorcraft; and total vertical load specified in sub paragraph (a)(1) is divided equally among the floats; and (ii) Distributed equally over 33.3% of the length between the skid (2) For each float, the load share tube attachments and centrally located determined under subpa ragraph (b)(1), midway between the skid tube combined with a total sideload of 0.25 times attachments.

the total vertical load specified in sub paragraph (b)(1), is applied to the float only.

CS 27.505 Ski landing conditions MAIN COMPONENT REQUIREMENTS If certification for ski operation is requested, the rotorcraft, with skis, must be designed to CS 27.547 Main rotor structure withstand the following loading conditions (where P is the maximum static weight on each (a) Each main rotor assembly (including ski with the rotorcraft at design maximum rotor hubs and blades) must be designed as weight, and n is the limit load factor determined prescribed in this paragraph.

under CS 27.473(b)).

(b) (Reserved) (a) Up load conditions in which: (c) The main rotor structure must be (1) A vertical load of Pn and a designed to withstand the following loads horizontal load of Pn/4 are simultaneously prescribed in CS 27.337 to 27.341: applied at the pedestal bearings; and (1) Critical flight loads.

(2) A vertical load of 1.33 P is applied (2) Limit loads occurring under normal at the pedestal bearings.

conditions of autorotation. For this condition, (b) A side load condition in which a side the rotor rpm must be selected to include the load of 0.35 Pn is applied at the pedestal bearings effects of altitude.

in a horizontal plane perpendicular to the (d) The main rotor structure must be centreline of the rotorcraft.

designed to withstand loads simulating: (c) A torque load condition in which a (1) For the rotor blades, hubs, and torque load of 1.33 P (in foot pounds) is applied flapping hinges, the impact force of each blade to the ski about the vertical axis through the against its stop during ground operation; and centreline of the pedestal bearings.

Amendment 2 1–C–8 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (2) Any other critical condition (a) The rotorcraft, although it may be damaged expected in normal operation. in emergency landing conditions on land or water, must be designed as prescribed in this paragraph to (e) The main rotor structure must be protect the occupants under those conditions.

designed to withstand the limit torque at any rotational speed, including zero. In addition: (b) The structure must be designed to give each occupant every reasonable chance of (1) The limit torque need not be escaping serious injury in a crash landing when: greater than the torque defined by a torque limiting device (where provided), and may not (1) Proper use is made of seats, belts, be less than the greater of: and other safety design provisions; (i) The maximum torque likely (2) The wheels are retracted (where to be transmitted to the rotor structure in applicable); and either direction; and (3) Each occupant and each item of (ii) The limit engine torque mass inside the cabin that could injure an specified in CS 27.361. occupant is restrained when subjected to the following ultimate inertial load factors relative (2) The limit torque must be to the surrounding structure: distributed to the rotor blades in a rational manner. (i) Upward – 4 g (ii) Forward – 16 g CS 27.549 Fuselage, landing gear, and (iii) Sideward – 8 g rotor pylon structures (iv) Downward – 20 g, after the (a) Each fuselage, landing gear, and rotor intended displacement of the seat device pylon structure must be designed as prescribed in (v) Rearward – 1.5 g this paragraph. Resultant rotor forces may be represented as a single force applied at the rotor (c) The supporting structure must be hub attachment point.

designed to restrain, under any ultimate inertial load up to those specified in this paragraph, any (b) Each structure must be designed to item of mass above and/or behind the crew and withstand: passenger compartment that could injure an (1) The critical loads prescribed in CS occupant if it came loose in an emergency 27.337 to 27.341; landing. Items of mass to be considered include, but are not limited to, rotors, transmissions, and engines. The items of mass must be restrained (2) The applicable ground loads for the following ultimate inertial load factors: prescribed in CS 27.235, 27.471 to 27.485, CS 27.493, 27.497, 27.501, 27.505, and 27.521; (1) Upward – 1.5 g and (2) Forward – 12 g (3) The loads prescribed in CS 27.547 (3) Sideward – 6 g (d)(2) and (e).

(4) Downward – 12 g (c) Auxiliary rotor thrust, and the balancing air and inertia loads occurring under accelerated (5) Rearward – 1.5 g flight conditions, must be considered.

(d) Any fuselage structure in the area of (d) Each engine mount and adjacent fuselage internal fuel tanks below the passenger floor structure must be designed to withstand the loads level must be designed to resist the following occuring under accelerated flight and landing ultimate inertial factors and loads and to protect conditions, including engine torque.

the fuel tanks from rupture when those loads are applied to that area: (1) Upward – 1.5 g (2) Forward – 4.0 g EMERGENCY LANDING CONDITIONS (3) Sideward – 2.0 g CS 27.561 General (4) Downward – 4.0 g Amendment 2 1–C–9 Annex to ED Decision 2008/009/R CS–27 BOOK 1 CS 27.562 Emergency landing dynamic must be misaligned with respect to each other by conditions at least 10° vertically (i.e. pitch out of parallel) and by at least a 10° lateral roll, with the (a) The rotorcraft, although it may be directions optional, to account for possible floor damaged in an emergency crash landing, must be warp.

designed to reasonably protect each occupant when: (c) Compliance with the following must be shown: (1) The occupant properly uses the seats, safety belts, and shoulder harnesses (1) The seating device system must provided in the design; and remain intact although it may experience separation intended as part of its design.

(2) The occupant is exposed to the loads resulting from the conditions prescribed (2) The attachment between the seating in this paragraph.

device and the airframe structure must remain intact, although the structure may have (b) Each seat type design or other seating exceeded its limit load.

device approved for crew or passenger occupancy during take off and landing must successfully (3) The ATD’s shoulder harness strap complete dynamic tests or be demonstrated by or straps must remain on or in the immediate rational analysis based on dynamic tests of a vicinity of the ATD’s shoulder during the similar type seat in accordance with the impact.

following criteria. The tests must be conducted (4) The safety belt must remain on the with an occupant, simulated by a 77 kg (170 ATD’s pelvis during the impact.

pound) anthropomorphic test dummy (ATD), sitting in the normal upright position.

(5) The ATD’s head either does not contact any portion of the crew or passenger (1) A change in downward velocity of compartment, or if contact is made, the head not less than 9.1 m/s (30 ft/s) when the seat or impact does not exceed a head injury criteria other seating device is oriented in its nominal (HIC) of 1000 as determined by this equation.

position with respect to the rotorcraft’s reference system, the rotorcraft’s longitudinal 2.5 t ˘ È axis is canted upward 60° with respect to the 1 ˙ Í

( )

a(t)dt t t HIC = impact velocity vector, and the rotorcraft’s 1 ˙ Í

2 Ú

t t 1 2 ˙ Í t lateral axis is perpendicular to a vertical plane ˚ Î containing the impact velocity vector and the rotorcraft’s longitudinal axis. Peak floor Where: a(t) is the resultant acceleration at the deceleration must occur in not more than centre of gravity of the head form expressed as a 0.031 seconds after impact and must reach a multiple of g (the acceleration of gravity) and t minimum of 30 g.

t is the time duration, in seconds, of major head (2) A change in forward velocity of not impact, not to exceed 0.05 seconds.

less than 12.8 m/s (42 ft/s) when the seat or other (6) Loads in individual upper torso seating device is oriented in its nominal position harness straps must not exceed 7784 N (1750 with respect to the rotorcraft’s reference system, lbs). If dual straps are used for retaining the the rotorcraft’s longitudinal axis is yawed 10° upper torso, the total harness strap loads must either right or left of the impact velocity vector not exceed 8896 N (2000 lbs).

(whichever would cause the greatest load on the shoulder harness), the rotorcraft’s lateral axis is (7) The maximum compressive load contained in a horizontal plane containing the measured between the pelvis and the lumbar impact velocity vector, and the rotorcraft’s column of the ATD must not exceed 6674 N vertical axis is perpendicular to a horizontal (1500 lbs).

plane containing the impact velocity vector.

(d) An alternate approach that achieves an Peak floor deceleration must occur in not more equivalent or greater level of occupant protection, than 0.071 seconds after impact and must reach a as required by this paragraph, must be minimum of 18.4 g.

substantiated on a rational basis.

(3) Where floor rails or floor or sidewall attachment devices are used to attach the seating CS 27.563 Structural ditching provisions devices to the airframe structure for the conditions of this paragraph, the rails or devices Amendment 2 1–C–10 Annex to ED Decision 2008/009/R CS–27 BOOK 1 If certification with ditching provisions is (20 knots) between the rotorcraft and the requested, structural strength for ditching must water. The vertical load may not be less than meet the requirements of this paragraph and CS the highest likely buoyancy load determined 27.801(e). under sub paragraph (b) (1).

(a) Forward speed landing conditions . The rotorcraft must initially contact the most critical FATIGUE EVALUATION wave for reasonably probable water conditions at forward velocities from zero up to 56 Km/h (30 CS 27.571 Fatigue evaluation of flight knots) in likely pitch, roll and yaw attitudes. The structure rotorcraft limit vertical descent velocity may not be less than 1.5 m (5 ft) per second relative to the (a) General . Each portion of the flight mean water surface. Rotor lift may be used to act structure (the flight structure includes rotors, through the centre of gravity throughout the rotor drive systems between the engines and the landing impact. This lift may not exceed two rotor hubs, controls, fuselage, landing gear, and thirds of the design maximum weight. A their related primary attachments) the failure of maximum forward velocity of less than 56 km/h which could be catastrophic, must be identified (30 knots) may be used in design if it can be and must be evaluated under subpa ragraph (b), (c), demonstrated that the forward velocity selected (d), or (e). The following apply to each fatigue would not be exceeded in a normal one engine evaluation: out touchdown.

(1) The procedure for the evaluation (b) Auxiliary or emergency float conditions: must be approved.

(1) Floats fixed or deployed before (2) The locations of probable failure initial water contact . In addition to the must be determined.

landing loads in sub paragraph (a), each (3) In flight measurement must be auxiliary or emergency float, or its support included in determining the following: and attaching structure in the airframe or fuselage, must be designed for the load (i) Loads or stresses in all developed by a fully immersed float unless it critical conditions throughout the range can be shown that full immersion is unlikely.

of limitations in CS 27.309, except that If full immersion is unlikely, the highest likely manoeuvring load factors need not float buoyancy load must be applied. The exceed the maximum values expected in highest likely buoyancy load must include operation.

consideration of a partially immersed float (ii) The effect of altitude upon creating restoring moments to compensate the these loads or stresses.

upsetting moments caused by sidewind, unsymmetrical rotorcraft loading, water wave (4) The loading spectra must be as action, rotorcraft inertia and probable severe as those expected in operation structural damage and leakage considered including, but not limited to, external cargo under CS 27.801(d). Maximum roll and pitch operations, if applicable, and ground air angles determined from compliance with CS ground cycles. The loading spectra must be 27.801(d) may be used, if significant, to based on loads or stresses determined under determine the extent of immersion of each sub paragraph (a)(3).

float. If the floats are deployed in flight, (b) Fatigue tolerance evaluation . It must be appropriate air loads derived from the flight shown that the fatigue tolerance of the structure limitations with the floats deployed shall be ensures that the probability of catastrophic used in substantiation of the floats and their fatigue failure is extremely remote without attachment to the rotorcraft. For this purpose, establishing replacement times, inspection the design airspeed for limit load is the float intervals or other procedures under paragraph deployed airspeed operating limit multiplied A27.4 of appendix A.

by 1.11.

(c) Replacement time evaluation . It must be (2) Floats deployed after initial water shown that the probability of catastrophic fatigue contact . Each float must be designed for full failure is extremely remote within a replacement or partial immersion prescribed in sub time furnished under paragraph A27.4 of paragraph (b)(1). In addition, each float must appendix A.

be designed for combined vertical and drag loads using a relative limit speed of 37 Km/h Amendment 2 1–C–11 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (d) Fail safe evaluation . The following apply to fail safe evaluation: (1) It must be shown that all partial failures will become readily detectable under inspection procedures furnished under paragraph A27.4 of appendix A.

(2) The interval between the time when any partial failure becomes readily detectable under sub paragraph (d)(1), and the time when any such failure is expected to reduce the remaining strength of the structure to limit or maximum attainable loads (whichever is less), must be determined.

(3) It must be shown that the interval determined under sub paragraph (d)(2) is long enough, in relation to the inspection intervals and related procedures furnished under paragraph A27.4 of appendix A, to provide a probability of detection great enough to ensure that the probability of catastrophic failure is extremely remote.

(e) Combination of replacement time and fail safe evaluations. A component may be evaluated under a combination of sub paragraphs (c) and (d) . For such component it must be shown that the probability of catastrophic failure is extremely remote with an approved combination of replacement time, inspection intervals, and related procedures furnished under paragraph A27.4 of appendix A.

Amendment 2 1–C–12 Annex to ED Decision 2008/009/R CS 27 BOOK 1 SUBPART D – DESIGN AND CONSTRUCTION GENERAL CS 27.607 Fasteners (a) Each removable bolt, screw, nut, pin, or other fastener whose loss could jeopardise the safe CS 27.601 Design operation of the rotorcraft must incorporate two (a) The rotorcraft may have no design features separate locking devices. The fastener and its or details that experience has shown to be hazardous locking devices may not be adversely affected by the or unreliable. environmental conditions associated with the particular installation.

(b) The suitability of each questionable design detail and part must be established by tests. (b) No self locking nut may be used on any bolt subject to rotation in operation unless a non friction locking device is used in addition to the self CS 27.602 Critical parts locking device.

(a) Critical part A critical part is a part, the failure of which could have a catastrophic effect CS 27.609 Protection of structure upon the rotorcraft, and for which critical characteristics have been identified which must be Each part of the structure must: controlled to ensure the required level of integrity.

(a) Be suitably protected against deterioration (b) If the type design includes critical parts, a or loss of strength in service due to any cause, critical parts list shall be established. Procedures including: shall be established to define the critical design (1) Weathering; characteristics, identify processes that affect those characteristics, and identify the design change and (2) Corrosion; and process change controls necessary for showing (3) Abrasion; and compliance with the quality assurance requirements of Part 21. (b) Have provisions for ventilation and drainage where necessary to prevent the accumulation of corrosive, flammable, or noxious CS 27.603 Materials fluids.

The suitability and durability of materials used for parts, the failure of which could adversely affect CS 27.610 Lightning and static electricity safety, must: protection (a) Be established on the basis of experience or (a) The rotorcraft must be protected against tests; catastrophic effects from lightning.

(b) Meet approved specifications that ensure (b) For metallic components, compliance with their having the strength and other properties sub paragraph (a) may be shown by: assumed in the design data; and (1) Electrically bonding the components (c) Take into account the effects of properly to the airframe; or environmental conditions, such as temperature and humidity, expected in service. (2) Designing the components so that a strike will not endanger the rotorcraft.

(c) For non metallic components, compliance CS 27.605 Fabrication methods with sub paragraph (a) may be shown by: (a) The methods of fabrication used must (1) Designing the components to produce consistently sound structures. If a minimise the effect of a strike; or fabrication process (such as gluing, spot welding, or heat treating) requires close control to reach this (2) Incorporating acceptable means of objective, the process must be performed according diverting the resulting electrical current so as not to an approved process specification. to endanger the rotorcraft.

(b) Each new aircraft fabrication method must (d) The electrical bonding and protection be substantiated by a test program. against lightning and static electricity must: (1) Minimise the accumulation of electrostatic charge; Amendment 2 1–D–1 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (2) Minimise the risk of electric shock to specimen of each individual item is tested before use crew, passengers, and service and maintenance and it is determined that the actual strength personnel using normal precautions; properties of that particular item will equal or exceed those used in design.

(3) Provide an electrical return path, under both normal and fault conditions, on rotorcraft having grounded electrical systems; CS 27.619 Special factors and (a) The special factors prescribed in CS 27.621 (4) Reduce to an acceptable level the to 27.625 apply to each part of the structure whose effects of lightning and static electricity on the strength is: functioning of essential electrical and electronic equipment. (1) Uncertain; (2) Likely to deteriorate in service before normal replacement; or CS 27.611 Inspection provisions (3) Subject to appreciable variability due There must be means to allow the close to: examination of each part that requires: (i) Uncertainties in manufacturing (a) Recurring inspection; processes; or (b) Adjustment for proper alignment and (ii) Uncertainties in inspection functioning; or methods.

(c) Lubrication.

(b) For each part to which CS 27.621 to 27.625 apply, the factor of safety prescribed in CS 27.303 must be multiplied by a special factor equal to: CS 27.613 Material strength properties (1) The applicable special factors and design values prescribed in CS 27.621 to 27.625; or (a) Material strength properties must be based (2) Any other factor great enough to on enough tests of material meeting specifications to ensure that the probability of the part being establish design values on a statistical basis.

understrength because of the uncertainties (b) Design values must be chosen to minimise specified in sub paragraph (a) is extremely the probability of structural failure due to material remote.

variability. Except as provided in sub paragraphs (d) and (e), compliance with this paragraph must be shown by selecting design values that assure CS 27.621 Casting factors material strength with the following probability: (a) General . The factors, tests, and inspections (1) Where applied loads are eventually specified in sub paragraphs (b) and (c) must be distributed through a single member within an applied in addition to those necessary to establish assembly, the failure of which would result in foundry quality control. The inspections must meet loss of structural integrity of the component, 99% approved specifications. Sub paragraphs (c) and (d) probability with 95% confidence; and apply to structural castings except castings that are pressure tested as parts of hydraulic or other fluid (2) For redundant structure, those in systems and do not support structural loads.

which the failure of individual elements would result in applied loads being safely distributed to (b) Bearing stresses and surfaces . The casting other load carrying members, 90% probability factors specified in sub paragraphs (c) and (d): with 95% confidence.

(1) Need not exceed 1.25 with respect to bearing stresses regardless of the method of (c) The strength, detail design, and fabrication inspection used; and of the structure must minimise the probability of disastrous fatigue failure, particularly at points of (2) Need not be used with respect to the stress concentration.

bearing surfaces of a part whose bearing factor is larger than the applicable casting factor.

(d) Material specifications must be those contained in documents accepted by the Agency.

(c) Critical castings . For each casting whose failure would preclude continued safe flight and (e) Other design values may be used if a selection of the material is made in which a Amendment 2 1–D–2 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (i) A casting factor of l.0 may be landing of the rotorcraft or result in serious injury to used; and any occupant, the following apply: (ii) The castings must be inspected (1) Each critical casting must – as provided in sub paragraph (d)(1) for (i) Have a casting factor of not less casting factors of l.25 to 1.50 and tested than 1.25; and under sub paragraph (c)(2).

(ii) Receive 100% inspection by visual, radiographic, and magnetic particle CS 27.623 Bearing factors (for ferromagnetic materials) or penetrant (for non ferromagnetic materials) (a) Except as provided in sub paragraph (b), inspection methods or approved equivalent each part that has clearance (free fit), and that is inspection methods.

subject to pounding or vibration, must have a (2) For each critical casting with a bearing factor large enough to provide for the effects casting factor less than 1.50, three sample of normal relative motion.

castings must be static tested and shown to meet (b) No bearing factor need be used on a part for – which any larger special factor is prescribed.

(i) The strength requirements of CS 27.305 at an ultimate load corresponding to a casting factor of 1.25; CS 27.625 Fitting factors and For each fitting (part or terminal used to join one (ii) The deformation requirements structural member to another) the following apply: of CS 27.305 at a load of 1.15 times the limit load.

(a) For each fitting whose strength is not proven by limit and ultimate load tests in which (d) Non critical castings. For each casting actual stress conditions are simulated in the fitting other than those specified in sub paragraph (c), the and surrounding structures, a fitting factor of at least following apply: 1.15 must be applied to each part of: (1) Except as provided in sub paragraphs (1) The fitting; (d)(2) and (3), the casting factors and corresponding inspections must meet the (2) The means of attachment; and following table: (3) The bearing on the joined members.

Casting factor Inspection (b) No fitting factor need be used: 2.0 or greater......... 100% visual (1) For joints made under approved Less than 2.0 greater 100% visual and practices and based on comprehensive test data than 1.5 magnetic particle (ferro magnetic materials), (such as continuous joints in metal plating, penetrant (non welded joints, and scarf joints in wood); and ferromagnetic materials), or approved equivalent (2) With respect to any bearing surface inspection methods.

for which a larger special factor is used.

1.25 through 1.50...... 100% visual, and magnetic particle (ferro (c) For each integral fitting, the part must be magnetic materials), treated as a fitting up to the point at which the penetrant nonfe rro paragraph properties become typical of the member.

magnetic materials), and radiographic or approved (d) Each seat, berth, litter, safety belt, and equivalent inspection methods harness attachment to the structure must be shown by analysis, tests, or both, to be able to withstand (2) The percentage of castings inspected the inertia forces prescribed in CS 27.561(b)(3) by nonvisual methods may be reduced below that multiplied by a fitting factor of 1.33.

specified in sub paragraph (d)(1) when an approved quality control procedure is established.

(3) For castings procured to a CS 27.629 Flutter specification that guarantees the mechanical Each aerodynamic surface of the rotorcraft must properties of the material in the casting and be free from flutter under each appropriate speed and provides for demonstration of these properties by power condition.

test of coupons cut from the castings on a sampling basis: Amendment 2 1–D–3 Annex to ED Decision 2008/009/R CS 27 BOOK 1 ROTORS CONTROL SYSTEMS CS 27.653 Pressure venting and drainage CS 27.671 General of rotor blades (a) Each control and control system must (a) For each rotor blade: operate with the ease, smoothness, and positiveness appropriate to its function.

(1) There must be means for venting the internal pressure of the blade; (b) Each element of each flight control system must be designed, or distinctively and permanently (2) Drainage holes must be provided for marked, to minimise the probability of any incorrect the blade; and assembly that could result in the malfunction of the (3) The blade must be designed to system.

prevent water from becoming trapped in it.

(b) Sub paragraphs (a)(1) and (2) do not apply CS 27.672 Stability augmentation, to sealed rotor blades capable of withstanding the automatic, and power operated maximum pressure differentials expected in service.

systems If the functioning of stability augmentation or CS 27.659 Mass balance other automatic or power operated systems is necessary to show compliance with the flight (a) The rotors and blades must be mass characteristics requirements of this CS–27, such balanced as necessary to – systems must comply with CS 27.671 and the (1) Prevent excessive vibration; and following: (2) Prevent flutter at any speed up to the (a) A warning which is clearly distinguishable maximum forward speed. to the pilot under expected flight conditions without requiring the pilot’s attention must be provided for (b) The structural integrity of the mass balance any failure in the stability augmentation system or in installation must be substantiated.

any other automatic or power operated system which could result in an unsafe condition if the pilot is unaware of the failure. Warning systems must not CS 27.661 Rotor blade clearance activate the control systems.

There must be enough clearance between the (b) The design of the stability augmentation rotor blades and other parts of the structure to system or of any other automatic or power operated prevent the blades from striking any part of the system must allow initial counteraction of failures structure during any operating condition.

without requiring exceptional pilot skill or strength by overriding the failure by movement of the flight controls in the normal sense and deactivating the CS 27.663 Ground resonance prevention failed system.

means (c) It must be shown that after any single (a) The reliability of the means for preventing failure of the stability augmentation system or any ground resonance must be shown either by analysis other automatic or power operated system: and tests, or reliable service experience, or by showing through analysis or tests that malfunction (1) The rotorcraft is safely controllable or failure of a single means will not cause ground when the failure or malfunction occurs at any resonance. speed or altitude within the approved operating limitations; (b) The probable range of variations, during service, of the damping action of the ground (2) The controllability and resonance prevention means must be established and manoeuvrability requirements of this CS–27 are must be investigated during the test required by CS met within a practical operational flight envelope 27.241. (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 permit continued safe flight and landing.

Amendment 2 1–D–4 Annex to ED Decision 2008/009/R CS 27 BOOK 1 CS 27.673 Primary flight control (1) The direction of the test loads produces the most severe loading in the control Primary flight controls are those used by the pilot system; and for immediate control of pitch, roll, yaw, and vertical motion of the rotorcraft. (2) Each fitting, pulley, and bracket used in attaching the system to the main structure is included.

CS 27.674 Interconnected controls (b) Compliance must be shown (by analyses or Each primary flight control system must provide individual load tests) with the special factor for safe flight and landing and operate independently requirements for control system joints subject to after a malfunction, failure, or jam of any auxiliary angular motion.

interconnected control.

CS 27.683 Operation tests CS 27.675 Stops It must be shown by operation tests that, when (a) Each control system must have stops that the controls are operated from the pilot compartment positively limit the range of motion of the pilot’s with the control system loaded to correspond with controls. loads specified for the system, the system is free from: (b) Each stop must be located in the system so that the range of travel of its control is not (a) Jamming; appreciably affected by: (b) Excessive friction; and (1) Wear; (c) Excessive deflection.

(2) Slackness; or (3) Take up adjustments.

CS 27.685 Control system details (c) Each stop must be able to withstand the (a) Each detail of each control system must be loads corresponding to the design conditions for the designed to prevent jamming, chafing, and system.

interference from cargo, passengers, loose objects or (d) For each main rotor blade: the freezing of moisture.

(1) Stops that are appropriate to the blade (b) There must be means in the cockpit to design must be provided to limit travel of the prevent the entry of foreign objects into places where blade about its hinge points; and they would jam the system.

(2) There must be means to keep the (c) There must be means to prevent the blade from hitting the droop stops during any slapping of cables or tubes against other parts.

operation other than starting and stopping the (d) Cable systems must be designed as follows: rotor.

(1) Cables, cable fittings, turnbuckles, splices and pulleys must be of an acceptable CS 27.679 Control system locks kind.

If there is a device to lock the control system with (2) The design of the cable systems must the rotorcraft on the ground or water, there must be prevent any hazardous change in cable tension means to: throughout the range of travel under any operating conditions and temperature variations.

(a) Give unmistakable warning to the pilot when the lock is engaged; and (3) No cable smaller than 2.4 mm (3/32 inch) diameter may be used in any primary (b) Prevent the lock from engaging in flight.

control system.

(4) Pulley kinds and sizes must CS 27.681 Limit load static tests correspond to the cables with which they are used.

(a) Compliance with the limit load requirements of this CS–27 must be shown by tests (5) Pulleys must have close fitting guards in which: to prevent the cables from being displaced or fouled.

Amendment 2 1–D–5 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (6) Pulleys must lie close enough to the (b) Each alternate system may be a duplicate plane passing through the cable to prevent the power portion or a manually operated mechanical cable from rubbing against the pulley flange. system. The power portion includes the power source (such as hydraulic pumps), and such items (7) No fairlead may cause a change in as valves, lines, and actuators.

cable direction of more than 3°.

(c) The failure of mechanical parts (such as (8) No clevis pin subject to load or piston rods and links), and the jamming of power motion and retained only by cotter pins may be cylinders, must be considered unless they are used in the control system.

extremely improbable.

(9) Turnbuckles attached to parts having angular motion must be installed to prevent binding throughout the range of travel.

LANDING GEAR (10) There must be means for visual inspection at each fairlead, pulley, terminal and turnbuckle. CS 27.723 Shock absorption tests (e) Control system joints subject to angular The landing inertia load factor and the reserve motion must incorporate the following special energy absorption capacity of the landing gear must factors with respect to the ultimate bearing strength be substantiated by the tests prescribed in CS 27.725 of the softest material used as a bearing: and 27.727, respectively. These tests must be conducted on the complete rotorcraft or on units (1) 3.33 for push pull systems other than consisting of wheel, tyre, and shock absorber in their ball and roller bearing systems.

proper relation.

(2) 2.0 for cable systems.

(f) For control system joints, the CS 27.725 Limit drop test manufacturer’s static, non Brinell rating of ball and roller bearings must not be exceeded.

The limit drop test must be conducted as follows: (a) The drop height must be – CS 27.687 Spring devices (1) 0.33 m (13 inches) from the lowest point of the landing gear to the ground; or (a) Each control system spring device where failure could cause flutter or other unsafe (2) Any lesser height, not less than characteristics must be reliable.

0.20 m (8 in), resulting in a drop contact velocity equal to the greatest probable sinking speed (b) Compliance with sub paragraph (a) must likely to occur at ground contact in normal be shown by tests simulating service conditions.

power off landings.

(b) If considered, the rotor lift specified in CS CS 27.691 Autorotation control 27.473(a) must be introduced into the drop test by mechanism appropriate energy absorbing devices or by the use of an effective mass.

Each main rotor blade pitch control mechanism must allow rapid entry into autorotation after power (c) Each landing gear unit must be tested in the failure.

attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it.

CS 27.695 Power boost and power (d) When an effective mass is used in showing operated control system compliance with sub paragraph (b) the following (a) If a power boost or power operated control formula may be used instead of more rational system is used, an alternate system must be computations: immediately available that allows continued safe W flight and landing in the event of: L)d (1 h - + e L n n and : W W + = = j e W d h + (1) Any single failure in the power portion of the system; or where: (2) The failure of all engines.

Amendment 2 1–D–6 Annex to ED Decision 2008/009/R CS 27 BOOK 1 W = the effective weight to be used in the drop (c) The landing gear must withstand this test e test .

without collapsing. Collapse of the landing gear occurs when a member of the nose, tail, or main gear W=W for main gear units , equal to the static M will not support the rotorcraft in the proper attitude reaction on the particular unit with the or allows the rotorcraft structure, other than the rotorcraft in the most critical attitude. A rational method may be used in computing landing gear and external accessories, to impact the a main gear static reaction, taking into landing surface.

consideration the moment arm between the main wheel reaction and the rotorcraft centre of gravity.

CS 27.729 Retracting mechanism W=W for nose gear units , equal to the vertical N For rotorcraft with retractable landing gear, the component of the static reaction that would following apply: exist at the nose wheel, assuming that the (a) Loads. The landing gear, retracting mass of the rotorcraft acts at the centre of mechanism, wheel well doors, and supporting gravity and exerts a force of 1.0 g structure must be designed for – downward and 0.25 g forward.

(1) The loads occurring in any W=W for tailwheel units equal to whichever of the T following is critical: manoeuvring condition with the gear retracted; (1) The static weight on the tailwheel (2) The combined friction, inertia, and air with the rotorcraft resting on all wheels; or loads occurring during retraction and extension at any airspeed up to the design maximum landing (2) The vertical component of the ground gear operating speed; and reaction that would occur at the tailwheel, assuming that the mass of the rotorcraft acts at (3) The flight loads, including those in the centre of gravity and exerts a force of 1 g yawed flight, occurring with the gear extended at downward with the rotorcraft in the maximum any airspeed up to the design maximum landing nose up attitude considered in the nose up gear extended speed.

landing conditions.

(b) Landing gear lock. A positive means must h = specified free drop height . be provided to keep the gear extended.

L = ratio of assumed rotor lift to the rotorcraft (c) Emergency operation. When other than weight. manual power is used to operate the gear, emergency means must be provided for extending the gear in d = deflection under impact of the tyre (at the the event of – proper inflation pressure) plus the vertical component of the axle travel relative to the (1) Any reasonably probable failure in drop mass. the normal retraction system; or n = limit inertia load factor. (2) The failure of any single source of hydraulic, electric, or equivalent energy.

n = the load factor developed, during impact, j on the mass used in the drop test (i.e., the (d) Operation tests . The proper functioning of acceleration dv/dt in g recorded in the drop the retracting mechanism must be shown by test plus 1.0). operation tests.

(e) Position indicator . There must be a means to indicate to the pilot when the gear is secured in CS 27.727 Reserve energy absorption the extreme positions.

drop test (f) Control . The location and the operation of The reserve energy absorption drop test must be the retraction control must meet the requirements of conducted as follows: CS 27.777 and 27.779.

(a) The drop height must be 1.5 times that (g) Landing gear warning . An aural or equally specified in CS 27.725(a).

effective landing gear warning device must be (b) Rotor lift, where considered in a manner provided that functions continuously when the similar to that prescribed in CS 27.725(b), may not rotorcraft is in a normal landing mode and the exceed 1.5 times the lift allowed under that landing gear is not fully extended and locked. A paragraph. manual shut off capability must be provided for the warning device and the warning system must Amendment 2 1–D–7 Annex to ED Decision 2008/009/R CS 27 BOOK 1 automatically reset when the rotorcraft is no longer CS 27.737 Skis in the landing mode.

The maximum limit load rating of each ski must equal or exceed the maximum limit load determined under the applicable ground load requirements of CS 27.731 Wheels this CS–27.

(a) Each landing gear wheel must be approved.

(b) The maximum static load rating of each wheel may not be less than the corresponding static FLOATS AND HULLS ground reaction with: (1) Maximum weight; and CS 27.751 Main float buoyancy (2) Critical centre of gravity.

(a) For main floats, the buoyancy necessary to (c) The maximum limit load rating of each support the maximum weight of the rotorcraft in wheel must equal or exceed the maximum radial fresh water must be exceeded by: limit load determined under the applicable ground (1) 50%, for single floats; and load requirements of this CS–27.

(2) 60%, for multiple floats.

(b) Each main float must have enough CS 27.733 Tyres watertight compartments so that, with any single (a) Each landing gear wheel must have a tyre: main float compartment flooded, the main floats will provide a margin of positive stability great enough (1) That is a proper fit on the rim of the to minimise the probability of capsizing.

wheel; and (2) Of the proper rating.

CS 27.753 Main float design (b) The maximum static load rating of each tyre must equal or exceed the static ground reaction (a) Bag floats . Each bag float must be obtained at its wheel, assuming: designed to withstand: (1) The design maximum weight; and (1) The maximum pressure differential that might be developed at the maximum altitude (2) The most unfavourable centre of for which certification with that float is gravity.

requested; and (c) Each tyre installed on a retractable landing (2) The vertical loads prescribed in CS gear system must, at the maximum size of the tyre 27.521(a), distributed along the length of the bag type expected in service, have a clearance to over three quarters of its projected area.

surrounding structure and systems that is adequate to prevent contact between the tyre and any part of (b) Rigid floats . Each rigid float must be able the structure or systems. to withstand the vertical, horizontal, and side loads prescribed in CS 27.521. These loads may be distributed along the length of the float.

CS 27.735 Brakes For rotorcraft with wheel type landing gear, a CS 27.755 Hulls braking device must be installed that is: For each rotorcraft, with a hull and auxiliary (a) Controllable by the pilot; floats, that is to be approved for both taking off from (b) Usable during power off landings; and and landing on water, the hull and auxiliary floats must have enough watertight compartments so that, (c) Adequate to: with any single compartment flooded, the buoyancy (1) Counteract any normal unbalanced of the hull and auxiliary floats (and wheel tyres if torque when starting or stopping the rotor; and used) provides a margin of positive stability great enough to minimise the probability of capsizing.

(2) Hold the rotorcraft parked on a 10° slope on a dry, smooth pavement.

Amendment 2 1–D–8 Annex to ED Decision 2008/009/R CS 27 BOOK 1 PERSONNEL AND CARGO CS 27.779 Motion and effect of cockpit ACCOMMODATIONS controls Cockpit controls must be designed so that they operate in accordance with the following movements CS 27.771 Pilot compartment and actuation: For each pilot compartment: (a) Flight controls, including the collective (a) The compartment and its equipment must pitch control, must operate with a sense of motion allow each pilot to perform his duties without which corresponds to the effect on the rotorcraft.

unreasonable concentration or fatigue; (b) Twist grip engine power controls must be (b) If there is provision for a second pilot, the designed so that, for left hand operation, the motion rotorcraft must be controllable with equal safety of the pilot’s hand is clockwise to increase power from either pilot seat; and when the hand is viewed from the edge containing the index finger. Other engine power controls, (c) The vibration and noise characteristics of excluding the collective control, must operate with a cockpit appurtenances may not interfere with safe forward motion to increase power.

operation.

(c) Normal landing gear controls must operate downward to extend the landing gear.

CS 27.773 Pilot compartment view (a) Each pilot compartment must be free from CS 27.783 Doors glare and reflections that could interfere with the pilot’s view, and designed so that: (a) Each closed cabin must have at least one adequate and easily accessible external door.

(1) Each pilot’s view is sufficiently extensive, clear, and undistorted for safe (b) Each external door must be located where operation; and persons using it will not be endangered by the rotors, propellers, engine intakes and exhausts when (2) Each pilot is protected from the appropriate operating procedures are used. If elements so that moderate rain conditions do not opening procedures are required, they must be the unduly impair his view of the flight path in marked inside, on or adjacent to the door opening normal flight and while landing.

device.

(b) If certification for night operation is requested, compliance with sub paragraph (a) must be shown in night flight tests. CS 27.785 Seats, berths, safety belts, and harnesses (a) Each seat, safety belt, harness, and adjacent CS 27.775 Windshields and windows part of the rotorcraft at each station designated for Windshields and windows must be made of occupancy during take off and landing must be free material that will not break into dangerous of potentially injurious objects, sharp edges, fragments. protuberances, and hard surfaces and must be designed so that a person making proper use of these facilities will not suffer serious injury in an CS 27.777 Cockpit controls emergency landing as a result of the static inertial load factors specified in CS 27.561(b) and dynamic Cockpit controls must be: conditions specified in CS 27.562.

(a) Located to provide convenient operation (b) Each occupant must be protected from and to prevent confusion and inadvertent operation; serious head injury by a safety belt plus a shoulder and harness that will prevent the head from contacting (b) Located and arranged with respect to the any injurious object except as provided for in CS pilots’ seats so that there is full and unrestricted 27.562(c)(5). A shoulder harness (upper torso movement of each control without interference from restraint), in combination with the safety belt, the cockpit structure or the pilot’s clothing when constitutes a torso restraint system as described in pilots from 1.57 m (5 ft 2 inches) to 1.83 m (6 ft) in ETSO C114.

height are seated.

(c) Each occupant’s seat must have a combined safety belt and shoulder harness with a single point release. Each pilot’s combined safety belt and Amendment 2 1–D–9 Annex to ED Decision 2008/009/R CS 27 BOOK 1 shoulder harness must allow each pilot when seated parts of the seats to reduce occupant loads for the with safety belt and shoulder harness fastened, to emergency landing dynamic conditions of CS perform all functions necessary for flight operations. 27.562; otherwise, the system must remain intact There must be a means to secure belts and harnesses and must not interfere with rapid evacuation of the when not in use, to prevent interference with the rotorcraft.

operation of the rotorcraft and with rapid egress in (k) For the purposes of this paragraph, a litter an emergency.

is defined as a device designed to carry a non (d) If seat backs do not have a firm handhold, ambulatory person, primarily in a recumbent there must be hand grips or rails along each aisle to position, into and on the rotorcraft. Each berth or enable the occupants to steady themselves while litter must be designed to withstand the load using the aisle in moderately rough air. reaction of an occupant weight of at least 77 kg (170 lbs) when the occupant is subjected to the forward (e) Each projecting object that could injure inertial factors specified in CS 27.561(b). A berth persons seated or moving about in the rotorcraft in or litter installed within 15° or less of the normal flight must be padded.

longitudinal axis of the rotorcraft must be provided (f) Each seat and its supporting structure must with a padded end board, cloth diaphragm, or designed for an occupant weight of at least 77 kg equivalent means that can withstand the forward (170 lbs) considering the maximum load factors, load reaction. A berth or litter oriented greater than inertial forces, and reactions between the occupant, 15° with the longitudinal axis of the rotorcraft must seat, and safety belt or harness corresponding with be equipped with appropriate restraints, such as the applicable flight and ground load conditions, straps or safety belts, to withstand the forward load including the emergency landing conditions of CS reaction. In addition – 27.561(b). In addition: (1) The berth or litter must have a (1) Each pilot seat must be designed for restraint system and must not have corners or the reactions resulting from the application of the other protuberances likely to cause serious injury pilot forces prescribed in CS 27.397; and to a person occupying it during emergency landing conditions; and (2) The inertial forces prescribed in CS 27.561(b) must be multiplied by a factor of 1.33 (2) The berth or litter attachment and the in determining the strength of the attachment of: occupant restraint system attachments to the structure must be designed to withstand the (i) Each seat to the structure; and critical loads resulting from flight and ground (ii) Each safety belt or harness to load conditions and from the conditions the seat or structure.

prescribed in CS 27.561(b). The fitting factor required by CS 27.625(d) shall be applied.

(g) When the safety belt and shoulder harness are combined, the rated strength of the safety belt and shoulder harness may not be less than that CS 27.787 Cargo and baggage corresponding to the inertial forces specified in CS compartments 27.561(b), considering the occupant weight of at least 77 kg (170 lbs), considering the dimensional (a) Each cargo and baggage compartment must characteristics of the restraint system installation, be designed for its placarded maximum weight of and using a distribution of at least a 60% load to the contents and for the critical load distributions at the safety belt and at least a 40% load to the shoulder appropriate maximum load factors corresponding to harness. If the safety belt is capable of being used the specified flight and ground load conditions, without the shoulder harness, the inertial forces except the emergency landing conditions of CS specified must be met by the safety belt alone.

27.561.

(h) When a headrest is used, the headrest and (b) There must be means to prevent the its supporting structure must be designed to resist contents of any compartment from becoming a the inertia forces specified in CS 27.561, with a 1.33 hazard by shifting under the loads specified in sub fitting factor and a head weight of at least 5.9 kg (13 paragraph (a).

lbs).

(c) Under the emergency landing conditions of (i) Each seating device system includes the CS 27.561, cargo and baggage compartments must: device such as the seat, the cushions, the occupant (1) Be positioned so that if the contents restraint system, and attachment devices.

break loose they are unlikely to cause injury to (j) Each seating device system may use design the occupants or restrict any of the escape features such as crushing or separation of certain Amendment 2 1–D–10 Annex to ED Decision 2008/009/R CS 27 BOOK 1 facilities provided for use after an emergency CS 27.805 Flight crew emergency exits landing; or (a) For rotorcraft with passenger emergency (2) Have sufficient strength to withstand exits that are not convenient to the flight crew, there the conditions specified in CS 27.561 including must be flight crew emergency exits, on both sides the means of restraint, and their attachments, of the rotorcraft or as a top hatch, in the flight crew required by sub paragraph (b). Sufficent strength area.

must be provided for the maximum authorised (b) Each flight crew emergency exit must be of weight of cargo and baggage at the critical sufficient size and must be located so as to allow loading distribution.

rapid evacuation of the flight crew. This must be (d) If cargo compartment lamps are installed, shown by test.

each lamp must be installed so as to prevent contact (c) Each flight crew emergency exit must not between lamp bulb and cargo.

be obstructed by water or flotation devices after an emergency landing on water. This must be shown by test, demonstration, or analysis.

CS 27.801 Ditching (a) If certification with ditching provisions is requested, the rotorcraft must meet the requirements CS 27.807 Emergency exits of this paragraph and CS 27.807(d), 27.1411 and (a) Number and location.

27.1415.

(1) There must be at least one emergency (b) Each practicable design measure, exit on each side of the cabin readily accessible to compatible with the general characteristics of the each passenger. One of these exits must be usable rotorcraft, must be taken to minimise the probability in any probable attitude that may result from a that in an emergency landing on water, the crash; behaviour of the rotorcraft would cause immediate injury to the occupants or would make it impossible (2) Doors intended for normal use may for them to escape. also serve as emergency exits, provided that they meet the requirements of this paragraph; and (c) The probable behaviour of the rotorcraft in a water landing must be investigated by model tests (3) If emergency flotation devices are or by comparison with rotorcraft of similar installed, there must be an emergency exit configuration for which the ditching characteristics accessible to each passenger on each side of the are known. Scoops, flaps, projections, and any other cabin that is shown by test, demonstration, or factor likely to affect the hydrodynamic analysis to: characteristics of the rotorcraft must be considered.

(i) Be above the waterline; and (d) It must be shown that, under reasonably (ii) Open without interference from probable water conditions, the flotation time and flotation devices, whether stowed or trim of the rotorcraft will allow the occupants to deployed.

leave the rotorcraft and enter the life rafts required (b) Type and operation . Each emergency exit by CS 27.1415. If compliance with this provision is prescribed by sub paragraph (a) must: shown by buoyancy and trim computations, appropriate allowances must be made for probable (1) Consist of a moveable window or structural damage and leakage. If the rotorcraft has panel, or additional external door, providing an fuel tanks (with fuel jettisoning provisions) that can unobstructed opening that will admit a 0.48 m by reasonably be expected to withstand a ditching 0.66 m (19 inch by 26 inch) ellipse; without leakage, the jettisonable volume of fuel may (2) Have simple and obvious methods of be considered as buoyancy volume.

opening, from the inside and from the outside, (e) Unless the effects of the collapse of external which do not require exceptional effort; doors and windows are accounted for in the (3) Be arranged and marked so as to be investigation of the probable behaviour of the readily located and operated even in darkness; rotorcraft in a water landing (as prescribed in sub and paragraphs (c) and (d)), the external doors and windows must be designed to withstand the (4) Be reasonably protected from probable maximum local pressures.

jamming by fuselage deformation.

(c) Tests. The proper functioning of each emergency exit must be shown by test.

Amendment 2 1–D–11 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (d) Ditching emergency exits for passengers. CS 27.855 Cargo and baggage If certification with ditching provisions is requested, compartments the markings required by sub paragraph (b)(3) must (a) Each cargo and baggage compartment must be designed to remain visible if the rotorcraft is be constructed of, or lined with, materials that are at capsized and the cabin is submerged.

least: (1) Flame resistant, in the case of CS 27.831 Ventilation compartments that are readily accessible to a crew member in flight; and (a) The ventilating system for the pilot and passenger compartments must be designed to (2) Fire resistant, in the case of other prevent the presence of excessive quantities of fuel compartments.

fumes and carbon monoxide.

(b) No compartment may contain any controls, (b) The concentration of carbon monoxide may wiring, lines, equipment, or accessories whose not exceed one part in 20 000 parts of air during damage or failure would affect safe operation, unless forward flight or hovering in still air. If the those items are protected so that: concentration exceeds this value under other (1) They cannot be damaged by the conditions, there must be suitable operating movement of cargo in the compartment; and restrictions.

(2) Their breakage or failure will not create a fire hazard.

CS 27.833 Heaters Each combustion heater must be approved.

CS 27.859 Heating systems (a) General. For each heating system that involves the passage of cabin air over, or close to, FIRE PROTECTION the exhaust manifold, there must be means to prevent carbon monoxide from entering any cabin or pilot compartment.

CS 27.853 Compartment interiors (b) Heat exchangers. Each heat exchanger For each compartment to be used by the crew or must be: passengers: (1) Of suitable materials; (a) The materials must be at least flame resistant; (2) Adequately cooled under all conditions; and (b) (Reserved) (3) Easily disassembled for inspection.

(c) If smoking is to be prohibited, there must be a placard so stating, and if smoking is to be (c) Combustion heater fire protection. Except allowed: for heaters which incorporate designs to prevent hazards in the event of fuel leakage in the heater fuel (1) There must be an adequate number of system, fire within the ventilating air passage, or self contained, removable ashtrays; and any other heater malfunction, each heater zone must (2) Where the crew compartment is incorporate the fire protection features of the separated from the passenger compartment, there applicable requirements of CS 27.1183, 27.1185, must be at least one illuminated sign (using 27.1189, 27.1191, and be provided with – either letters or symbols) notifying all passengers (1) Approved, quick acting fire detectors when smoking is prohibited. Signs which notify in numbers and locations ensuring prompt when smoking is prohibited must: detection of fire in the heater region.

(i) When illuminated, be legible to (2) Fire extinguisher systems that provide each passenger seated in the passenger at least one adequate discharge to all areas of the cabin under all probable lighting heater region.

conditions; and (3) Complete drainage of each part of (ii) Be so constructed that the crew can turn the illumination on and off. each zone to minimise the hazards resulting from failure or malfunction of any component containing flammable fluids. The drainage means must be: Amendment 2 1–D–12 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (i) Effective under conditions (i) The heat exchanger temperature expected to prevail when drainage is exceeds safe limits.

needed; and (ii) The ventilating air temperature (ii) Arranged so that no discharged exceeds safe limits.

fluid will cause an additional fire hazard.

(iii) The combustion airflow becomes inadequate for safe operation.

(4) Ventilation, arranged so that no discharged vapours will cause an additional fire (iv) The ventilating airflow becomes hazard.

inadequate for safe operation.

(d) Ventilating air ducts . Each ventilating air (2) The means of complying with sub duct passing through any heater region must be paragraph (g)(1) for any individual heater must: fireproof.

(i) Be independent of components (1) Unless isolation is provided by fire serving any other heater, the heat output of proof valves or by equally effective means, the which is essential for safe operation; and ventilating air duct downstream of each heater (ii) Keep the heater off until must be fireproof for a distance great enough to restarted by the crew.

ensure that any fire originating in the heater can (3) There must be means to warn the be contained in the duct.

crew when any heater, the heat output of which is (2) Each part of any ventilating duct essential for safe operation, has been shut off by passing through any region having a flammable the automatic means prescribed in sub paragraph fluid system must be so constructed or isolated (g)(1).

from that system that the malfunctioning of any (h) Air intakes. Each combustion and heat component of that system cannot introduce ventilating air intake must be located so that no flammable fluids or vapours into the ventilating flammable fluids or vapours can enter the heater airstream.

system: (e) Combustion air ducts. Each combustion (1) During normal operation; or air duct must be fireproof for a distance great enough to prevent damage from backfiring or reverse (2) As a result of the malfunction of any flame propagation.

other component.

(1) No combustion air duct may connect (i) Heater exhaust . Each heater exhaust with the ventilating airstream unless flames from system must meet the requirements of CS 27.1121 back fires or reverse burning cannot enter the and 27.1123.

ventilating airstream under any operating (1) Each exhaust shroud must be sealed condition, including reverse flow or malfunction so that no flammable fluids or hazardous of the heater or its associated components.

quantities of vapours can reach the exhaust (2) No combustion air duct may restrict system through joints.

the prompt relief of any backfire that, if so (2) No exhaust system may restrict the restricted, could cause heater failure.

prompt relief of any backfire that, if so restricted, could cause heater failure.

(f) Heater control. General. There must be means to prevent the hazardous accumulation of (j) Heater fuel systems. Each heater fuel water or ice on or in any heater control component, system must meet the powerplant fuel system control system tubing, or safety control.

requirements affecting safe heater operation. Each heater fuel system component in the ventilating (g) Heater safety controls. For each airstream must be protected by shrouds so that no combustion heater, safety control means must be leakage from those components can enter the provided as follows: ventilating airstream.

(1) Means independent of the (k) Drains. There must be means for safe components provided for the normal continuous drainage of any fuel that might accumulate in the control of air temperature, airflow, and fuel flow combustion chamber or the heat exchanger.

must be provided for each heater to automatically shut off the ignition and fuel supply of that heater (1) Each part of any drain that operates at at a point remote from that heater when any of the high temperatures must be protected in the same following occurs: manner as heater exhausts.

Amendment 2 1–D–13 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (2) Each drain must be protected against EXTERNAL LOADS hazardous ice accumulation under any operating condition.

CS 27.865 External loads (a) It must be shown by analysis, test, or both, CS 27.861 Fire protection of structure, that the rotorcraft external load attaching means for controls, and other parts rotorcraft load combinations to be used for non Each part of the structure, controls, rotor human external cargo applications can withstand a mechanism, and other parts essential to a controlled limit static load equal to 2.5, or some lower load landing that would be affected by powerplant fires factor approved under CS 27.337 through 27.341, must be fireproof or protected so they can perform multiplied by the maximum external load for which their essential functions for at least 5 minutes under authorisation is requested. It must be shown by any foreseeable powerplant fire conditions. analysis, test, or both that the rotorcraft external load attaching means and corresponding personnel carrying device system for rotorcraft load CS 27.863 Flammable fluid fire protection combinations to be used for human external cargo applications can withstand a limit static load equal (a) In each area where flammable fluids or to 3.5 or some lower load factor, not less than 2.5, vapours might escape by leakage of a fluid system, approved under CS 27.337 through 27.341, there must be means to minimise the probability of multiplied by the maximum external load for which ignition of the fluids and vapours, and the resultant authorisation is requested. The load for any hazards if ignition does occur.

rotorcraft load combination class, for any external (b) Compliance with sub paragraph (a) must cargo type, must be applied in the vertical direction.

be shown by analysis or tests, and the following For jettisonable rotorcraft load combinations, for any factors must be considered: applicable external cargo type, the load must also be applied in any direction making the maximum angle (1) Possible sources and paths of fluid with the vertical that can be achieved in service but leakage, and means of detecting leakage.

not less than 30º. However, the 30º angle may be (2) Flammability characteristics of fluids, reduced to a lesser angle if: including effects of any combustible or absorbing (1) An operating limitation is established materials.

limiting external load operations to such angles (3) Possible ignition sources, including for which compliance with this paragraph has electrical faults, over heating of equipment, and been shown; or malfunctioning of protective devices.

(2) It is shown that the lesser angle can (4) Means available for controlling or not be exceeded in service.

extinguishing a fire, such as stopping flow of (b) The external load attaching means, for fluids, shutting down equipment, fireproof jettisonable rotorcraft load combinations, must containment, or use of extinguishing agents.

include a quick release system to enable the pilot to (5) Ability of rotorcraft components that release the external load quickly during flight. The are critical to safety of flight to withstand fire and quick release system must consist of a primary heat. quick release subsystem and a backup quick release subsystem that are isolated from one another. The (c) If action by the flight crew is required to quick release system, and the means by which it is prevent or counteract a fluid fire (e.g. equipment controlled, must comply with the following: shutdown or actuation of a fire extinguisher) quick acting means must be provided to alert the crew. (1) A control for the primary quick release subsystem must be installed either on one (d) Each area where flammable fluids or of the pilot's primary controls or in an vapours might escape by leakage of a fluid system equivalently accessible location and must be must be identified and defined.

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, must be provided.

Amendment 2 1–D–14 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (3) Both the primary and backup quick (5) Have the appropriate limitations and release subsystems must: procedures incorporated in the flight manual for conducting human external cargo operations.

(i) Be reliable, durable, and function properly with all external loads up (6) For human external cargo to and including the maximum external applications requiring use of Category A limit load for which authorisation is rotorcraft, have one engine inoperative hover requested.

performance data and procedures in the flight (ii) Be protected against manual for the weights, altitudes, and electromagnetic interference (EMI) from temperatures for which external load approval is external and internal sources and against requested.

lightning to prevent inadvertent load (d) The critically configured jettisonable release.

external loads must be shown by a combination of (A) The minimum level of analysis, ground tests, and flight tests to be both protection required for jettisonable transportable and releasable throughout the rotorcraft load combinations used for approved operational envelope without hazard to the non human external cargo is a radio rotorcraft during normal flight conditions. In frequency field strength of 20 volts addition, these external loads must be shown to be per metre.

releasable without hazard to the rotorcraft during emergency flight conditions.

(B) The minimum level of protection required for jettisonable (e) A placard or marking must be installed rotorcraft load combinations used for next to the external load attaching means stating the human external cargo is a radio maximum authorised external load as demonstrated frequency field strength of 200 volts under CS 27.25 and this paragraph.

per metre.

(f) The fatigue evaluation of CS 27.571(a) (iii) Be protected against any failure does not apply to this paragraph except for a failure that could be induced by a failure mode of of the cargo attaching means that results in a hazard any other electrical or mechanical rotorcraft to the rotorcraft.

system.

(c) For rotorcraft load combinations to be used for human external cargo applications, the rotorcraft MISCELLANEOUS must: (1) For jettisonable external loads, have a quick release system that meets the requirements CS 27.871 Levelling marks of sub paragraph (b) and that: There must be reference marks for levelling the (i) Provides a dual actuation rotorcraft on the ground.

device for the primary quick release subsystem, and (ii) Provides a separate dual CS 27.873 Ballast provisions actuation device for the backup quick Ballast provisions must be designed and release subsystem.

constructed to prevent inadvertent shifting of ballast (2) Have a reliable, approved in flight.

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, the ingress and egress instructions.

(4) Have equipment to allow direct intercommunication among required crew members and external occupants, and Amendment 2 1–D–15 Annex to ED Decision 2008/009/R CS 27 BOOK 1 SUBPART E – POWERPLANT GENERAL CS 27.903 Engines CS 27.901 Installation (a) (Reserved) (a) For the purpose of this CS–27, the (b) Engine or drive system cooling fan powerplant installation includes each part of the blade protection.

rotorcraft (other than the main and auxiliary (1) If an engine or rotor drive system rotor structures) that: cooling fan is installed, there must be means (1) Is necessary for propulsion; to protect the rotorcraft and allow a safe landing if a fan blade fails. This must be (2) Affects the control of the major shown by showing that: propulsive units; or (i) The fan blades are contained (3) Affects the safety of the major in case of failure; propulsive units between normal inspections or overhauls.

(ii) Each fan is located so that a failure will not jeopardise safety; or (b) For each powerplant installation: (iii) Each fan blade can (1) Each component of the withstand an ultimate load of 1.5 times installation must be constructed, arranged, the centrifugal force resulting from and installed to ensure its continued safe operation limited by the following: operation between normal inspections or overhauls for the range of temperature and (A) For fans driven altitude for which approval is requested; directly by the engine: (2) Accessibility must be provided to (1) The terminal allow any inspection and maintenance engine rpm under necessary for continued airworthiness; uncontrolled conditions; or (3) Electrical interconnections must (2) An overspeed be provided to prevent differences of limiting device.

potential between major components of the (B) For fans driven by the installation and the rest of the rotorcraft; rotor drive system the maximum (4) Axial and radial expansion of rotor drive system rotational turbine engines may not affect the safety of speed to be expected in service, the installation; and including transients.

(5) Design precautions must be taken (2) Unless a fatigue evaluation under to minimise the possibility of incorrect CS 27.571 is conducted, it must be shown assembly of components and equipment that cooling fan blades are not operating at essential to safe operation of the rotorcraft, resonant conditions within the operating except where operation with the incorrect limits of the rotorcraft.

assembly can be shown to be extremely (c) Turbine engine installation. For turbine improbable.

engine installations, the powerplant systems (c) The installation must comply with: associated with engine control devices, systems, and instrumentation must be designed to give (1) The installation instructions reasonable assurance that those engine operating provided under CS–E; and limitations that adversely affect turbine rotor (2) The applicable provisions of this structural integrity will not be exceeded in service.

Subpart.

(d) Restart capability : A means to restart any engine in flight must be provided.

Amendment 2 1–E–1 Annex to ED Decision 2008/009/R CS 27 BOOK 1 CS 27.921 Rotor brake (1) Except for the in flight shutdown of all engines, engine restart capability must If there is a means to control the rotation of be demonstrated throughout a flight envelope the rotor drive system independently of the for the rotorcraft.

engine, any limitations on the use of that means must be specified, and the control for that (2) Following the in flight shutdown means must be guarded to prevent inadvertent of all engines, in flight engine restart operation.

capability must be provided.

[Amdt. No.: 27/1] CS 27.923 Rotor drive system and control mechanism tests CS 27.907 Engine vibration (a) Each part tested as prescribed in this (a) Each engine must be installed to prevent paragraph must be in a serviceable condition at the harmful vibration of any part of the engine or the end of the tests. No intervening rotorcraft.

disassembly which might affect test results may be conducted.

(b) The addition of the rotor and the rotor drive system to the engine may not subject the (b) Each rotor drive system and control principal rotating parts of the engine to excessive mechanism must be tested for not less than vibration stresses. This must be shown by a 100 hours. The test must be conducted on the vibration investigation.

rotorcraft, and the torque must be absorbed by the rotors to be installed, except that other (c) No part of the rotor drive system may ground or flight test facilities with other be subjected to excessive vibration stresses.

appropriate methods of torque absorption may be used if the conditions of support and vibration closely simulate the conditions that ROTOR DRIVE SYSTEM would exist during a test on the rotorcraft.

(c) A 60 hour part of the test prescribed in sub paragraph (b) must be run at not less than CS 27.917 Design maximum continuous torque and the maximum (a) Each rotor drive system must speed for use with maximum continuous torque.

incorporate a unit for each engine to In this test, the main rotor controls must be set automatically disengage that engine from the in the position that will give maximum main and auxiliary rotors if that engine fails. longitudinal cyclic pitch change to simulate forward flight. The auxiliary rotor controls (b) Each rotor drive system must be must be in the position for normal operation arranged so that each rotor necessary for control under the conditions of the test.

in autorotation will continue to be driven by the main rotors after disengagement of the engine (d) A 30 hour or, for rotorcraft for which from the main and auxiliary rotors. the use of either 30 minute OEI power or continuous OEI power is requested, a 25 hour (c) If a torque limiting device is used in part of the test prescribed in sub paragraph (b) the rotor drive system, it must be located so as must be run at not less than 75% of maximum to allow continued control of the rotorcraft continuous torque and the minimum speed for when the device is operating.

use with 75% of maximum continuous torque.

(d) The rotor drive system includes any The main and auxiliary rotor controls must be part necessary to transmit power from the in the position for normal operation under the engines to the rotor hubs. This includes gear conditions of the test.

boxes, shafting, universal joints, couplings, (e) A 10 hour part of the test prescribed in rotor brake assemblies, clutches, supporting sub paragraph (b) must be run at not less than bearings for shafting, any attendant accessory take off torque and the maximum speed for use pads or drives, and any cooling fans that are a with take off torque. The main and auxiliary part of, attached to, or mounted on the rotor rotor controls must be in the normal position for drive system.

vertical ascent.

(1) For multi engine rotorcraft for which the use of 2½ minute OEI power is Amendment 2 1–E–2 Annex to ED Decision 2008/009/R CS 27 BOOK 1 requested, 12 runs during the 10 hour test show compliance with the remainder of must be conducted as follows: this paragraph.

(i) Each run must consist of at (f) The parts of the test prescribed in sub least one period of 2½ minutes with paragraphs (c) and (d) must be conducted in take off torque and the maximum speed intervals of not less than 30 minutes and may be for use with take off torque on all accomplished either on the ground or in flight.

engines.

The part of the test prescribed in sub paragraph (e) must be conducted in intervals of not less (ii) Each run must consist of at than 5 minutes.

least one period for each engine in sequence, during which that engine (g) At intervals of not more than five hours simulates a power failure and the during the tests prescribed in sub paragraphs remaining engines are run at 2½ (c), (d), and (e), the engine must be stopped minute OEI torque and the maximum rapidly enough to allow the engine and rotor speed for use with 2½ minute OEI drive to be automatically disengaged from the torque for 2½ minutes.

rotors.

(2) For multi engine turbine powered (h) Under the operating conditions rotorcraft for which the use of 30 second and specified in sub paragraph (c), 500 complete 2 minute OEI power is requested, 10 runs cycles of lateral control, 500 complete cycles of must be conducted as follows: longitudinal control of the main rotors, and 500 (i) Immediately following a complete cycles of control of each auxiliary take off run of at least 5 minutes, each rotor must be accomplished. A ‘complete cycle’ power source must simulate a failure, involves movement of the controls from the neutral position, through both extreme in turn, and apply the maximum torque positions, and back to the neutral position, and the maximum speed for use with except that control movements need not produce 30 second OEI power to the remaining loads or flapping motions exceeding the affected drive system power inputs for maximum loads or motions encountered in not less than 30 seconds, followed by flight. The cycling may be accomplished during application of the maximum torque and the testing prescribed in sub paragraph (c).

the maximum speed for use with 2 minute OEI power for not less than (i) At least 200 start up clutch 2 minutes. At least one run sequence engagements must be accomplished: must be conducted from a simulated (1) So that the shaft on the driven ‘flight idle’ condition. When side of the clutch is accelerated; and conducted on a bench test, the test sequence must be conducted following (2) Using a speed and method stabilisation at take off power.

selected by the applicant.

(ii) For the purpose of this (j) For multi engine rotorcraft for which paragraph, an affected power input the use of 30 minute OEI power is requested, includes all parts of the rotor drive five runs must be made at 30 minute OEI torque system which can be adversely affected and the maximum speed for use with 30 minute by the application of higher or OEI torque, in which each engine, in sequence, asymmetric torque and speed prescribed is made inoperative and the remaining by the test.

engine(s) is run for a 30 minute period.

(iii) This test may be conducted (k) For multi engine rotorcraft for which on a representative bench test facility the use of continuous OEI power is requested, when engine limitations either preclude five runs must be made at continuous OEI repeated use of this power or would torque and the maximum speed for use with result in premature engine removal continuous OEI torque, in which each engine, during the test. The loads, the in sequence, is made inoperative and the vibration frequency, and the methods of remaining engine(s) is run for a 1 hour period.

application to the affected rotor drive system components must be representative of rotorcraft conditions.

Test components must be those used to CS 27.927 Additional tests Amendment 2 1–E–3 Annex to ED Decision 2008/009/R CS 27 BOOK 1 methods of analysis are available for the (a) Any additional dynamic, endurance, and operational tests, and vibratory particular design.

investigations necessary to determine that the (b) If any critical speed lies within, or rotor drive mechanism is safe, must be close to, the operating ranges for idling, power performed.

on, and autorotative conditions, the stresses occurring at that speed must be within safe (b) If turbine engine torque output to the limits. This must be shown by tests.

transmission can exceed the highest engine or transmission torque rating limit, and that output (c) If analytical methods are used and show is not directly controlled by the pilot under that no critical speed lies within the permissible normal operating conditions (such as where the operating ranges, the margins between the primary engine power control is accomplished calculated critical speeds and the limits of the through the flight control), the following test allowable operating ranges must be adequate to allow for possible variations between the must be made: computed and actual values.

(1) Under conditions associated with all engines operating, make 200 applications, for 10 seconds each, of torque that is at least equal to the lesser of: CS 27.935 Shafting joints (i) The maximum torque used in Each universal joint, slip joint, and other meeting CS 27.923 plus 10%; or shafting joints whose lubrication is necessary (ii) The maximum attainable for operation must have provision for torque output of the engines, assuming lubrication.

that torque limiting devices, if any, function properly.

(2) For multi engine rotorcraft under conditions associated with each engine in CS 27.939 Turbine engine operating turn becoming inoperative, apply to the characteristics remaining transmission torque inputs, the (a) Turbine engine operating maximum torque attainable under probable characteristics must be investigated in flight to operating conditions, assuming that torque determine that no adverse characteristics (such limiting devices, if any, function properly.

as stall, surge, or flameout) are present, to a Each transmission input must be tested at this maximum torque for at least 15 minutes. hazardous degree, during normal and emergency operation within the range of (3) The tests prescribed in this operating limitations of the rotorcraft and of the paragraph must be conducted on the engine.

rotorcraft at the maximum rotational speed intended for the power condition of the test (b) The turbine engine air inlet system may and the torque must be absorbed by the rotors not, as a result of airflow distortion during to be installed, except that other ground or normal operation, cause vibration harmful to flight test facilities with other appropriate the engine.

methods of torque absorption may be used if the conditions of support and vibration (c) For governor controlled engines, it closely simulate the conditions that would must be shown that there exists no hazardous exist during a test on the rotorcraft.

torsional instability of the drive system associated with critical combinations of power, (c) It must be shown by tests that the rotor rotational speed, and control displacement.

drive system is capable of operating under autorotative conditions for 15 minutes after the loss of pressure in the rotor drive primary oil system.

FUEL SYSTEM CS 27.931 Shafting critical speed CS 27.951 General (a) The critical speeds of any shafting must (a) Each fuel system must be constructed be determined by demonstration except that and arranged to ensure a flow of fuel at a rate analytical methods may be used if reliable and pressure established for proper engine functioning under any likely operating Amendment 2 1–E–4 Annex to ED Decision 2008/009/R CS 27 BOOK 1 condition, including the manoeuvres for which (b) Fuel tank load factors. Except for fuel certification is requested. tanks located so that tank rupture with fuel release to either significant ignition sources, (b) Each fuel system must be arranged so such as engines, heaters, and auxiliary power that: units, or occupants is extremely remote, each (1) No fuel pump can draw fuel from fuel tank must be designed and installed to more than one tank at a time; or retain its contents under the following ultimate inertial load factors, acting alone.

(2) There are means to prevent introducing air into the system. (1) For fuel tanks in the cabin: (c) Each fuel system for a turbine engine (i) Upward – 4 g.

must be capable of sustained operation (ii) Forward – 16 g.

throughout its flow and pressure range with fuel initially saturated with water at 27°C (80°F) (iii) Sideward – 8 g.

and having 0.198 cc of free water per litre (0.75 (iv) Downward – 20 g.

cc per US gallon) added and cooled to the most critical condition for icing likely to be (2) For fuel tanks located above or encountered in operation. behind the crew or passenger compartment that, if loosened, could injure an occupant in an emergency landing: (i) Upward – 1.5 g.

CS 27.952 Fuel system crash resistance (ii) Forward – 8 g.

Unless other means acceptable to the Agency are employed to minimise the hazard of fuel (iii) Sideward – 2 g.

fires to occupants following an otherwise (iv) Downward – 4 g.

survivable impact (crash landing), the fuel systems must incorporate the design features of (3) For fuel tanks in other areas: this paragraph. These systems must be shown (i) Upward – 1.5 g.

to be capable of sustaining the static and dynamic deceleration loads of this paragraph, (ii) Forward – 4 g.

considered as ultimate loads acting alone, (iii) Sideward – 2 g.

measured at the system component’s centre of gravity without structural damage to system (iv) Downward – 4 g.

components, fuel tanks, or their attachments ( c) Fuel line self sealing breakaway that would leak fuel to an ignition source.

couplings. Self sealing breakaway couplings (a) Drop test requirements. Each tank, or must be installed unless hazardous relative the most critical tank, must be drop tested as motion of fuel system components to each other follows: or to local rotorcraft structure is demonstrated to be extremely improbable or unless other (1) The drop height must be at least means are provided. The couplings or 15.2 m (50 ft).

equivalent devices must be installed at all fuel (2) The drop impact surface must be tank to fuel line connections, tank to tank non deforming.

interconnects, and at other points in the fuel system where local structural deformation could (3) The tank must be filled with lead to release of fuel.

water to 80% of the normal, full capacity.

(1) The design and construction of (4) The tank must be enclosed in a self sealing breakaway couplings must surrounding structure representative of the incorporate the following design features: installation unless it can be established that the surrounding structure is free of (i) The load necessary to separate projections or other design features likely to a breakaway coupling must be between contribute to rupture of the tank.

25 and 50% of the minimum ultimate failure load (ultimate strength) of the (5) The tank must drop freely and weakest component in the fluid impact in a horizontal position ±10°.

carrying line. The separation load must (6) After the drop test there must be in no case be less than 1334 N (300 lb), no leakage.

regardless of the size of the fluid line.

Amendment 2 1–E–5 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (ii) A breakaway coupling must (1) The load required to separate a separate whenever its ultimate load (as frangible attachment from its support defined in sub paragraph (c)(1)(i)) is structure, or to deform a locally deformable applied in the failure modes most likely attachment relative to its support structure, to occur. must be between 25 and 50% of the minimum ultimate load (ultimate strength) of (iii) All breakaway couplings the weakest component in the attached must incorporate design provisions to system. In no case may the load be less than visually ascertain that the coupling is 1330 N (300 lbs).

locked together (leak free) and is open during normal installation and service. (2) A frangible or locally deformable attachment must separate or locally deform (iv) All breakaway couplings as intended whenever its ultimate load (as must incorporate design provisions to defined in sub paragraph (d) (1)) is applied prevent uncoupling or unintended in the modes most likely to occur.

closing due to operational shocks, vibrations, or accelerations. (3) All frangible or locally deformable attachments must comply with (v) No breakaway coupling the fatigue requirements of CS 27.571.

design may allow the release of fuel once the coupling has performed its (e) Separation of fuel and ignition sources.

intended function. To provide maximum crash resistance, fuel must be located as far as practicable from all (2) All individual breakaway occupiable areas and from all potential ignition couplings, coupling fuel feed systems, or sources.

equivalent means must be designed, tested, installed and maintained so that inadvertent (f) Other basic mechanical design criteria.

Fuel tanks, fuel lines, electrical wires, and fuel shut off in flight is improbable in electrical devices must be designed, constructed accordance with CS 27.955(a) and must and installed, as far as practicable, to be crash comply with the fatigue evaluation resistant.

requirements of CS 27.571 without leaking.

(g) Rigid or semi rigid fuel tanks. Rigid or (3) Alternate, equivalent means to the semi rigid fuel tank or bladder walls must be use of breakaway couplings must not create a impact and tear resistant.

survivable impact induced load on the fuel line to which it is installed greater than 25 to 50% of the ultimate load (strength) of the weakest component of the line and must CS 27.953 Fuel system independence comply with the fatigue requirements of CS 27.571 without leaking.

(a) Each fuel system for multi engine rotorcraft must allow fuel to be supplied to each (d) Frangible or deformable structural engine through a system independent of those attachments. Unless hazardous relative motion parts of each system supplying fuel to other of fuel tanks and fuel system components to engines. However, separate fuel tanks need not local rotorcraft structure is demonstrated to be be provided for each engine.

extremely improbable in an otherwise (b) If a single fuel tank is used on a multi survivable impact, frangible or locally engine rotorcraft, the following must be deformable attachments of fuel tanks and fuel provided: system components to local rotorcraft structure must be used. The attachment of fuel tanks and (1) Independent tank outlets for each fuel system components to local rotorcraft engine, each incorporating a shut off valve at structure, whether frangible or locally the tank. This shut off valve may also serve deformable, must be designed such that its as the firewall shut off valve required by CS separation or relative local deformation will 27.995 if the line between the valve and the occur without rupture or local tear out of the engine compartment does not contain a fuel tank and fuel system components that will hazardous amount of fuel that can drain into cause fuel leakage. The ultimate strength of the engine compartment.

frangible or deformable attachments must be as (2) At least two vents arranged to follows: minimise the probability of both vents becoming obstructed simultaneously.

Amendment 2 1–E–6 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (3) Filler caps designed to minimise (5) Critical values of engine rotation the probability of incorrect installation or in speed, electrical power, or other sources of flight loss. fuel pump motive power must be applied.

(4) A fuel system in which those (6) Critical values of fuel properties parts of the system from each tank outlet to which adversely affect fuel flow must be any engine are independent of each part of applied.

each system supplying fuel to other engines.

(7) The fuel filter required by CS 27.997 must be blocked to the degree necessary to simulate the accumulation of fuel contamination required to activate the CS 27.954 Fuel system lightning indicator required by CS 27.1305(q).

protection (b) Fuel transfer systems. If normal The fuel system must be designed and operation of the fuel system requires fuel to be arranged to prevent the ignition of fuel vapour transferred to an engine feed tank, the transfer within the system by: must occur automatically via a system which has been shown to maintain the fuel level in the (a) Direct lightning strikes to areas having engine feed tank within acceptable limits during a high probability of stroke attachment; flight or surface operation of the rotorcraft.

(b) Swept lightning strokes to areas where swept strokes are highly probable; or (c) Multiple fuel tanks. If an engine can be supplied with fuel from more than one tank, the (c) Corona and streamering at fuel vent fuel systems must, in addition to having outlets.

appropriate manual switching capability, be designed to prevent interruption of fuel flow to that engine, without attention by the flightcrew, when any tank supplying fuel to that engine is CS 27.955 Fuel flow depleted of usable fuel during normal operation, and any other tank that normally supplies fuel (a) General . The fuel system for each to the engine alone contains usable fuel.

engine must be shown to provide the engine with at least 100% of the fuel required under each operating and manoeuvring condition to be approved for the rotorcraft including, as CS 27.959 Unusable fuel supply applicable, the fuel required to operate the engine(s) under the test conditions required by The unusable fuel supply for each tank must CS 27.927. Unless equivalent methods are used, be established as not less than the quantity at compliance must be shown by test during which which the first evidence of malfunction occurs the following provisions are met except that under the most adverse fuel feed condition combinations of conditions which are shown to occurring under any intended operations and be improbable need not be considered: flight manoeuvres involving that tank.

(1) The fuel pressure, corrected for critical accelerations, must be within the limits specified by the engine type certificate data sheet.

CS 27.961 Fuel system hot weather operation (2) The fuel level in the tank may not exceed that established as unusable fuel Each suction lift fuel system and other fuel supply for the tank under CS 27.959, plus the systems with features conducive to vapour minimum additional fuel necessary to formation must be shown by test to operate conduct the test. satisfactorily (within certification limits) when using fuel at a temperature of 43°C (110°F) (3) The fuel head between the tank under critical operating conditions including, if outlet and the engine inlet must be critical applicable, the engine operating conditions with respect to rotorcraft flight attitudes.

defined by CS 27.927 (b)(1) and (b)(2).

(4) The critical fuel pump (for pump fed systems) is installed to produce (by actual or simulated failure) the critical restriction to fuel flow to be expected from pump failure.

CS 27.963 Fuel tanks: general Amendment 2 1–E–7 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (a) Each fuel tank must be able to (a) Each fuel tank must be able to withstand, without failure, the vibration, withstand the applicable pressure tests in this inertia, fluid, and structural loads to which it paragraph without failure or leakage. If may be subjected in operation.

practicable, test pressures may be applied in a manner simulating the pressure distribution in (b) Each fuel tank of 38 litres (8.3 Imperial service.

gallons/10 US gallons) or greater capacity must (b) Each conventional metal tank, non have internal baffles, or must have external metallic tank with walls that are not supported support to resist surging.

by the rotorcraft structure, and integral tank (c) Each fuel tank must be separated from must be subjected to a pressure of 24 kPa (3.5 the engine compartment by a firewall. At least psi) unless the pressure developed during one half inch of clear airspace must be provided maximum limit acceleration or emergency between the tank and the firewall.

deceleration with a full tank exceeds this value, in which case a hydrostatic head, or equivalent test, (d) Spaces adjacent to the surfaces of fuel must be applied to duplicate the acceleration loads tanks must be ventilated so that fumes cannot as far as possible. However, the pressure need not accumulate in the tank compartment in case of exceed 24 kPa (3.5 psi) on surfaces not exposed to leakage. If two or more tanks have the acceleration loading.

interconnected outlets, they must be considered as one tank, and the airspaces in those tanks (c) Each non metallic tank with walls must be interconnected to prevent the flow of supported by the rotorcraft structure must be fuel from one tank to another as a result of a subjected to the following tests: difference in pressure between those airspaces.

(1) A pressure test of at least 14 kPa (e) The maximum exposed surface (2.0 psi). This test may be conducted on the temperature of any component in the fuel tank tank alone in conjunction with the test must be less, by a safe margin, than the lowest specified in sub paragraph (c)(2) .

expected auto ignition temperature of the fuel (2) A pressure test, with the tank or fuel vapour in the tank. Compliance with this mounted in the rotorcraft structure, equal to requirement must be shown under all operating the load developed by the reaction of the conditions and under all failure or malfunction contents, with the tank full, during maximum conditions of all components inside the tank.

limit acceleration or emergency deceleration.

(f) Each fuel tank installed in personnel However, the pressure need not exceed compartments must be isolated by fume proof 14 kPa (2.0 psi) on surfaces not exposed to and fuel proof enclosures that are drained and the acceleration loading.

vented to the exterior of the rotorcraft. The (d) Each tank with large unsupported or design and construction of the enclosures must unstiffened flat areas, or with other features provide necessary protection for the tank, must whose failure or deformation could cause be crash resistant during a survivable impact in leakage, must be subjected to the following test accordance with CS 27.952 and must be or its equivalent: adequate to withstand loads and abrasions to be expected in personnel compartments.

(1) Each complete tank assembly and its support must be vibration tested while (g) Each flexible fuel tank bladder or liner mounted to simulate the actual installation.

must be approved or shown to be suitable for the particular application and must be puncture (2) The tank assembly must be resistant. Puncture resistance must be shown by vibrated for 25 hours while two thirds full of meeting the ETSO C80, paragraph 16.0, any suitable fluid. The amplitude of vibration requirements using a minimum puncture force may not be less than 0.8 mm (1/32 inch), of 1646 N (370 lbs).

unless otherwise substantiated.

(h) Each integral fuel tank must have (3) The test frequency of vibration provisions for inspection and repair of its must be as follows: interior.

(i) If no frequency of vibration resulting from any rpm within the normal operating range of engine or rotor system speeds is critical, the test CS 27.965 Fuel tank tests frequency of vibration, in number of cycles per minute must, unless a Amendment 2 1–E–8 Annex to ED Decision 2008/009/R CS 27 BOOK 1 frequency based on a more rational projections that could cause wear of the liner calculation is used, be the number unless: obtained by averaging the maximum (i) There are means for and minimum power on engine speeds protection of the liner at those points; (rpm) for reciprocating engine powered or rotorcraft or 2000 rpm for turbine engine powered rotorcraft. (ii) The construction of the liner itself provides such protection.

(ii) If only one frequency of vibration resulting from any rpm within (b) Any spaces adjacent to tank surfaces the normal operating range of engine or must be adequately ventilated to avoid rotor system speeds is critical, that accumulation of fuel or fumes in those spaces frequency of vibration must be the test due to minor leakage. If the tank is in a sealed frequency. compartment, ventilation may be limited to drain holes that prevent clogging and excessive (iii) If more than one frequency pressure resulting from altitude changes. If of vibration resulting from any rpm flexible tank liners are installed, the venting within the normal operating range of arrangement for the spaces between the liner engine or rotor system speeds is and its container must maintain the proper critical, the most critical of these relationship to tank vent pressures for any frequencies must be the test frequency.

expected flight condition.

(4) Under sub paragraphs (d)(3)(ii) (c) The location of each tank must meet and (iii), the time of test must be adjusted to the requirements of CS 27.1185 (a) and (c).

accomplish the same number of vibration cycles as would be accomplished in 25 hours (d) No rotorcraft skin immediately adjacent at the frequency specified in sub paragraph to a major air outlet from the engine (d)(3)(i) . compartment may act as the wall of the integral tank.

(5) During the test, the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute through an angle of 15° on both sides of the horizontal (30° CS 27.969 Fuel tank expansion space total), about the most critical axis, for 25 hours. If motion about more than one axis is Each fuel tank or each group of fuel tanks likely to be critical, the tank must be rocked with interconnected vent systems must have an about each critical axis for 12½ hours. expansion space of not less than 2% of the tank capacity. It must be impossible to fill the fuel tank expansion space inadvertently with the rotorcraft in the normal ground attitude.

CS 27.967 Fuel tank installation (a) Each fuel tank must be supported so that tank loads are not concentrated on CS 27.971 Fuel tank sump unsupported tank surfaces. In addition: (a) Each fuel tank must have a drainable (1) There must be pads, if necessary, sump with an effective capacity in any ground to prevent chafing between each tank and its attitude to be expected in service of 0.25% of supports; the tank capacity or 0.24 litres (0.05 Imperial gallons/one sixteenth US gallon), whichever is (2) The padding must be non greater, unless: absorbent or treated to prevent the absorption of fuel; (1) The fuel system has a sediment bowl or chamber that is accessible for pre (3) If flexible tank liners are used, flight drainage and has a minimum capacity they must be supported so that it is not of 30 ml (l ounce) for every 76 litres necessary for them to withstand fluid loads; (16.7 Imperial gallons/20 US gallons) of fuel and tank capacity; and (4) Each interior surface of tank (2) Each fuel tank drain is located so compartments must be smooth and free of that in any ground attitude to be expected in Amendment 2 1–E–9 Annex to ED Decision 2008/009/R CS 27 BOOK 1 service, water will drain from all parts of the (2) For turbine engine powered tank to the sediment bowl or chamber. rotorcraft, prevent the passage of any object that could restrict fuel flow or damage any (b) Each sump, sediment bowl, and fuel system component.

sediment chamber drain required by the paragraph must comply with the drain (b) The clear area of each fuel tank outlet provisions of CS 27.999 (b). strainer must be at least 5 times the area of the outlet line.

(c) The diameter of each strainer must be at least that of the fuel tank outlet.

CS 27.973 Fuel tank filler connection (d) Each finger strainer must be accessible (a) Each fuel tank filler connection must for inspection and cleaning.

prevent the entrance of fuel into any part of the rotorcraft other than the tank itself during normal operations and must be crash resistant FUEL SYSTEM COMPONENTS during a survivable impact in accordance with CS 27.952 (c). In addition: (1) Each filler must be marked as prescribed in CS 27.1557 (c)(1); CS 27.991 Fuel pumps (2) Each recessed filler connection Compliance with CS 27.955 may not be that can retain any appreciable quantity of jeopardised by failure of: fuel must have a drain that discharges clear (a) Any one pump except pumps that are of the entire rotorcraft; and approved and installed as parts of a type (3) Each filler cap must provide a certificated engine; or fuel tight seal under the fluid pressure (b) Any component required for pump expected in normal operation and in a operation except, for engine driven pumps, the survivable impact.

engine served by that pump.

(b) Each filler cap or filler cap cover must warn when the cap is not fully locked or seated on the filler connection.

CS 27.993 Fuel system lines and fittings (a) Each fuel line must be installed and supported to prevent excessive vibration and to CS 27.975 Fuel tank vents withstand loads due to fuel pressure and (a) Each fuel tank must be vented from the accelerated flight conditions.

top part of the expansion space so that venting (b) Each fuel line connected to components is effective under all normal flight conditions.

of the rotorcraft between which relative motion Each vent must minimise the probability of could exist must have provisions for flexibility.

stoppage by dirt or ice.

(c) Flexible hose must be approved.

(b) The venting system must be designed to minimise spillage of fuel through the vents to (d) Each flexible connection in fuel lines an ignition source in the event of a rollover that may be under pressure or subjected to axial during landing, ground operation, or a loading must use flexible hose assemblies.

survivable impact.

(e) No flexible hose that might be adversely affected by high temperatures may be used where excessive temperatures will exist during operation or after engine shutdown.

CS 27.977 Fuel tank outlet (a) There must be a fuel strainer for the fuel tank outlet or for the booster pump. This strainer must: CS 27.995 Fuel valves (1) For reciprocating engine powered (a) There must be a positive, quick acting rotorcraft have 3 to 6 meshes per cm (8 to 16 valve to shut off fuel to each engine meshes per inch); and individually.

Amendment 2 1–E–10 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (b) The control for this valve must be (3) Have a drain valve: within easy reach of appropriate crew members.

(i) That is readily accessible (c) Where there is more than one source of and which can be easily opened and fuel supply there must be means for closed; and independent feeding from each source.

(ii) That is either located or (d) No shut off valve may be on the engine protected to prevent fuel spillage in the side of any firewall. event of a landing with landing gear retracted.

CS 27.997 Fuel strainer or filter OIL SYSTEM There must be a fuel strainer or filter between the fuel tank outlet and the inlet of the first fuel system component which is susceptible CS 27.1011 Engines: general to fuel contamination, including but not limited to the fuel metering device or an engine positive (a) Each engine must have an independent displacement pump, whichever is nearer the oil system that can supply it with an appropriate fuel tank outlet. This fuel strainer or filter quantity of oil at a temperature not above that must: safe for continuous operation.

(a) Be accessible for draining and cleaning (b) The usable oil capacity of each system and must incorporate a screen or element which may not be less than the product of the is easily removable; endurance of the rotorcraft under critical operating conditions and the maximum oil (b) Have a sediment trap and drain except consumption of the engine under the same that it need not have a drain if the strainer or conditions, plus a suitable margin to ensure filter is easily removable for drain purposes; adequate circulation and cooling. Instead of a (c) Be mounted so that its weight is not rational analysis of endurance and consumption, supported by the connecting lines or by the inlet a usable oil capacity of 3.8 litres (0.83 Imperial or outlet connections of the strainer or filter gallon /l US gallon) for each 151 litres (33.3 itself, unless adequate strength margins under Imperial gallons/40 US gallons) of usable fuel all loading conditions are provided in the lines may be used.

and connections; and (c) The oil cooling provisions for each (d) Provide a means to remove from the engine must be able to maintain the oil inlet fuel any contaminant which would jeopardise temperature to that engine at or below the the flow of fuel through rotorcraft or engine fuel maximum established value. This must be system components required for proper shown by flight tests.

rotorcraft fuel system or engine fuel system operation.

CS 27.1013 Oil tanks Each oil tank must be designed and installed CS 27.999 Fuel system drains so that: (a) There must be at least one accessible (a) It can withstand, without failure, each drain at the lowest point in each fuel system to vibration, inertia, fluid, and structural load completely drain the system with the rotorcraft expected in operation; in any ground attitude to be expected in service.

(b) (Reserved) (b) Each drain required by sub paragraph (a) must: (c) Where used with a reciprocating engine, it has an expansion space of not less (1) Discharge clear of all parts of the than the greater of 10% of the tank capacity or rotorcraft; 1.9 litre (0.42 Imperial gallon/0.5 US gallon), (2) Have manual or automatic means and where used with a turbine engine, it has an to assure positive closure in the off position; expansion space of not less than 10% of the and tank capacity.

Amendment 2 1–E–11 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (d) It is impossible to fill the tank density) that is greater than that established expansion space inadvertently with the for the engine under CS–E.

rotorcraft in the normal ground attitude; (3) The oil strainer or filter, unless it (e) Adequate venting is provided; and is installed at an oil tank outlet, must incorporate a means to indicate (f) There are means in the filler opening to contamination before it reaches the capacity prevent oil overflow from entering the oil tank established in accordance with sub paragraph compartment.

(a)(2).

(4) The bypass of a strainer or filter must be constructed and installed so that the CS 27.1015 Oil tank tests release of collected contaminants is minimised by appropriate location of the Each oil tank must be designed and installed bypass to ensure that collected contaminants so that it can withstand, without leakage, an are not in the bypass flow path.

internal pressure of 34 kPa (5 psi), except that each pressurised oil tank used with a turbine (5) An oil strainer or filter that has engine must be designed and installed so that it no bypass, except one that is installed at an can withstand, without leakage, an internal oil tank outlet, must have a means to connect pressure of 34 kPa (5 psi), plus the maximum it to the warning system required in CS operating pressure of the tank. 27.1305(r).

(b) Each oil strainer or filter in a powerplant installation using reciprocating engines must be constructed and installed so CS 27.1017 Oil lines and fittings that oil will flow at the normal rate through the (a) Each oil line must be supported to rest of the system with the strainer or filter prevent excessive vibration.

element completely blocked.

(b) Each oil line connected to components of the rotorcraft between which relative motion could exist must have provisions for flexibility.

CS 27.1021 Oil system drains (c) Flexible hose must be approved.

A drain (or drains) must be provided to allow (d) Each oil line must have an inside safe drainage of the oil system. Each drain diameter of not less than the inside diameter of must: the engine inlet or outlet. No line may have (a) Be accessible; and splices between connections.

(b) Have manual or automatic means for positive locking in the closed position.

CS 27.1019 Oil strainer or filter (a) Each turbine engine installation must CS 27.1027 Transmissions and incorporate an oil strainer or filter through gearboxes: general which all of the engine oil flows and which meets the following requirements: (a) The lubrication system for components of the rotor drive system that require (1) Each oil strainer or filter that has continuous lubrication must be sufficiently a bypass must be constructed and installed so independent of the lubrication systems of the that oil will flow at the normal rate through engine(s) to ensure lubrication during the rest of the system with the strainer or autorotation.

filter completely blocked.

(b) Pressure lubrication systems for (2) The oil strainer or filter must transmissions and gear boxes must comply with have the capacity (with respect to operating the engine oil system requirements of CS limitations established for the engine) to 27.1013 (except sub paragraph (c)), CS ensure that engine oil system functioning is 27.1015, 27.1017, 27.1021, and 27.1337 (d).

not impaired when the oil is contaminated to a degree (with respect to particle size and Amendment 2 1–E–12 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (c) Each pressure lubrication system must cooling or lubricating fluids used with these have an oil strainer or filter through which all components.

of the lubricant flows and must: (b) Compliance with sub paragraph (a) (1) Be designed to remove from the must be shown in tests conducted under the lubricant any contaminant which may conditions prescribed in that paragraph.

damage transmission and drive system components or impede the flow of lubricant to a hazardous degree; CS 27.1043 Cooling tests (2) Be equipped with a means to indicate collection of contaminants on the (a) General . For the tests prescribed in CS filter or strainer at or before opening of the 27.1041 (b), the following apply: bypass required by sub paragraph (c)(3); and (1) If the tests are conducted under (3) Be equipped with a bypass conditions deviating from the maximum constructed and installed so that: ambient atmospheric temperature specified in sub paragraph (b), the recorded powerplant (i) The lubricant will flow at temperatures must be corrected under sub the normal rate through the rest of the paragraphs (c) and (d) unless a more rational system with the strainer or filter correction method is applicable.

completely blocked; and (2) No corrected temperature (ii) The release of collected determined under sub paragraph (a)(1) may contaminants is minimised by exceed established limits.

appropriate location of the bypass to ensure that collected contaminants are (3) For reciprocating engines, the not in the bypass flow path. fuel used during the cooling tests must be of the minimum grade approved for the (d) For each lubricant tank or sump outlet engines, and the mixture settings must be supplying lubrication to rotor drive systems and those normally used in the flight stages for rotor drive system components, a screen must be which the cooling tests are conducted.

provided to prevent entrance into the lubrication system of any object that might (4) The test procedures must be as obstruct the flow of lubricant from the outlet to prescribed in CS 27.1045.

the filter required by sub paragraph (c). The (b) Maximum ambient atmospheric requirements of sub paragraph (c) do not apply temperature. A maximum ambient atmospheric to screens installed at lubricant tank or sump temperature corresponding to sea level outlets.

conditions of at least 38°C (100°F) must be (e) Splash type lubrication systems for established. The assumed temperature lapse rate rotor drive system gearboxes must comply with is 1.98°C (3.6°F) per 305 m (1000 ft) of altitude CS 27.1021 and 27.1337 (d). above sea level until a temperature of 56.5°C ( 69.7°F) is reached, above which altitude the temperature is considered constant at 56.5°C ( 69.7°F). However, for winterization COOLING installations, the applicant may select a maximum ambient atmospheric temperature corresponding to sea level conditions of less than 38°C (100°F).

CS 27.1041 General (c) Correction factor (except cylinder (a) Each powerplant cooling system must barrels). Unless a more rational correction be able to maintain the temperatures of applies, temperatures of engine fluids and powerplant components within the limits powerplant components (except cylinder established for these components under critical barrels) for which temperature limits are surface (ground or water) and flight operating established, must be corrected by adding to conditions for which certification is required them the difference between the maximum and after normal shutdown. Powerplant ambient atmospheric temperature and the components to be considered include but may temperature of the ambient air at the time of the not be limited to engines, rotor drive system first occurrence of the maximum component or components, auxiliary power units, and the Amendment 2 1–E–13 Annex to ED Decision 2008/009/R CS 27 BOOK 1 fluid temperature recorded during the cooling (3) An operating limitation is test. reached.

(d) Correction factor for cylinder barrel temperatures. Cylinder barrel temperatures INDUCTION SYSTEM must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first CS 27.1091 Air induction occurrence of the maximum cylinder barrel temperature recorded during the cooling test. (a) The air induction system for each engine must supply the air required by that engine under the operating conditions and manoeuvres for which certification is requested.

CS 27.1045 Cooling test procedures (b) Each cold air induction system opening (a) General . For each stage of flight, the must be outside the cowling if backfire flames cooling tests must be conducted with the can emerge.

rotorcraft: (c) If fuel can accumulate in any air (1) In the configuration most critical induction system, that system must have drains for cooling; and that discharge fuel: (2) Under the conditions most critical (1) Clear of the rotorcraft; and for cooling.

(2) Out of the path of exhaust flames.

(b) Temperature stabilisation. For the (d) For turbine engine powered rotorcraft: purpose of the cooling tests, a temperature is ‘stabilised’ when its rate of change is less than 1°C (1) There must be means to prevent (2°F) per minute. The following component and hazardous quantities of fuel leakage or engine fluid temperature stabilisation rules apply: overflow from drains, vents, or other (1) For each rotorcraft, and for each components of flammable fluid systems from stage of flight: entering the engine intake system; and (i) The temperatures must be (2) The air inlet ducts must be stabilised under the conditions from located or protected so as to minimise the which entry is made into the stage of ingestion of foreign matter during take off, flight being investigated; or landing, and taxying.

(ii) If the entry condition normally does not allow temperatures to stabilise, operation through the full entry condition must be conducted CS 27.1093 Induction system icing before entry into the stage of flight protection being investigated in order to allow the (a) Reciprocating engines. Each temperatures to attain their natural reciprocating engine air induction system must levels at the time of entry.

have means to prevent and eliminate icing.

(2) For each helicopter during the Unless this is done by other means, it must be take off stage of flight the climb at take off shown that, in air free of visible moisture at a power must be preceded by a period of hover temperature of –1°C (30°F) and with the during which the temperatures are stabilised.

engines at 75% of maximum continuous power: (c) Duration of test. For each stage of (1) Each rotorcraft with sea level flight the tests must be continued until: engines using conventional venturi carburettors has a preheater that can provide (1) The temperatures stabilise or a heat rise of 50°C (90°F); 5 minutes after the occurrence of the highest temperature recorded, as appropriate to the (2) Each rotorcraft with sea level test condition; engines using carburettors tending to prevent icing has a sheltered alternate source of air, (2) That stage of flight is completed; and that the preheat supplied to the alternate or air intake is not less than that provided by Amendment 2 1–E–14 Annex to ED Decision 2008/009/R CS 27 BOOK 1 the engine cooling air downstream of the applicable altitude and operating condition cylinders; because of supercharging.

(3) Each rotorcraft with altitude engines using conventional venturi EXHAUST SYSTEM carburettors has a preheater capable of providing a heat rise of 67°C (120°F); and (4) Each rotorcraft with altitude CS 27.1121 General engines using carburettors tending to prevent icing has a preheater that can provide a heat For each exhaust system: rise of: (a) There must be means for thermal (i) 56°C (100°F); or expansion of manifolds and pipes; (ii) If a fluid de icing system is (b) There must be means to prevent local used, at least 22°C (40°F). hot spots; (b) Turbine engines (c) Exhaust gases must discharge clear of the engine air intake, fuel system components, (1) It must be shown that each and drains; turbine engine and its air inlet system can operate throughout the flight power range of (d) Each exhaust system part with a surface the engine (including idling): hot enough to ignite flammable fluids or vapours must be located or shielded so that (i) Without accumulating ice leakage from any system carrying flammable on engine or inlet system components fluids or vapours will not result in a fire caused that would adversely affect engine by impingement of the fluids or vapours on any operation or cause a serious loss of part of the exhaust system including shields for power under the icing conditions the exhaust system; specified in appendix C of CS–29; and (e) Exhaust gases may not impair pilot (ii) In snow, both falling and vision at night due to glare; blowing, without adverse effect on engine operation, within the limitations (f) If significant traps exist, each turbine established for the rotorcraft. engine exhaust system must have drains discharging clear of the rotorcraft, in any (2) Each turbine engine must idle for normal ground and flight attitudes, to prevent 30 minutes on the ground, with the air bleed fuel accumulation after the failure of an available for engine icing protection at its attempted engine start; and critical condition, without adverse effect, in an atmosphere that is at a temperature (g) Each exhaust heat exchanger must between –9°C and –1°C (15° and 30°F) and incorporate means to prevent blockage of the has a liquid water content not less than 0.3 exhaust port after any internal heat exchanger grams per cubic metre in the form of drops failure.

having a mean effective diameter of not less than 20 microns, followed by momentary operation at take off power or thrust. During the 30 minutes of idle operation, the engine CS 27.1123 Exhaust piping may be run up periodically to a moderate (a) Exhaust piping must be heat and power or thrust setting in a manner corrosion resistant and must have provisions to acceptable to the Agency.

prevent failure due to expansion by operating temperatures.

(c) Supercharged reciprocating engines.

For each engine having superchargers to (b) Exhaust piping must be supported to pressurise the air before it enters the withstand any vibration and inertia loads to carburettor, the heat rise in the air caused by which it would be subjected in operations.

that supercharging at any altitude may be (c) Exhaust piping connected to utilised in determining compliance with sub components between which relative motion paragraph (a) if the heat rise utilised is that could exist must have provisions for flexibility.

which will be available, automatically, for the Amendment 2 1–E–15 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (d) If a power control incorporates a fuel POWERPLANT CONTROLS AND shut off feature, the control must have a means ACCESSORIES to prevent the inadvertent movement of the control into the shut off position. The means must: (1) Have a positive lock or stop at the CS 27.1141 Powerplant controls: general idle position; and (a) Powerplant controls must be located (2) Require a separate and distinct and arranged under CS 27.777 and marked operation to place the control in the shut off under CS 27.1555. position.

(b) Each flexible powerplant control must (e) For rotorcraft to be certificated for a be approved. 30 second OEI power rating, a means must be provided to automatically activate and control (c) Each control must be able to maintain the 30 second OEI power and prevent any any set position without: engine from exceeding the installed engine (1) Constant attention; or limits associated with the 30 second OEI power rating approved for the rotorcraft.

(2) Tendency to creep due to control loads or vibration.

(d) Controls of powerplant valves required for safety must have: CS 27.1145 Ignition switches (1) For manual valves, positive stops (a) There must be means to quickly shut or in the case of fuel valves suitable index off all ignition by the grouping of switches or by provisions, in the open and closed position; a master ignition control.

and (b) Each group of ignition switches, except (2) For power assisted valves, a ignition switches for turbine engines for which means to indicate to the flight crew when the continuous ignition is not required, and each valve: master ignition control must have a means to prevent its inadvertent operation.

(i) Is in the fully open or fully closed position; or (ii) Is moving between the fully open and fully closed position. CS 27.1147 Mixture controls If there are mixture controls, each engine (e) For turbine engine powered rotorcraft, must have a separate control and the controls no single failure or malfunction, or probable must be arranged to allow: combination thereof, in any powerplant control system may cause the failure of any powerplant (a) Separate control of each engine; and function necessary for safety.

(b) Simultaneous control of all engines.

CS 27.1143 Engine controls CS 27.1151 Rotor brake controls (a) There must be a separate power control for each engine. (a) It must be impossible to apply the rotor brake inadvertently in flight.

(b) Power controls must be grouped and arranged to allow: (b) There must be means to warn the crew if the rotor brake has not been completely (1) Separate control of each engine; released before take off.

and (2) Simultaneous control of all engines.

CS 27.1163 Powerplant accessories (c) Each power control must provide a positive and immediately responsive means of (a) Each engine mounted accessory must: controlling its engine.

Amendment 2 1–E–16 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (1) Be approved for mounting on the (b) Each tank or reservoir, other than a engine involved; fuel tank, that is part of a system containing flammable fluids or gases must be isolated from (2) Use the provisions on the engine the engine by a firewall or shroud unless the for mounting; and design of the system, the materials used in the (3) Be sealed in such a way as to tank and its supports, the shutoff means, and prevent contamination of the engine oil the connections, lines and controls provide a system and the accessory system. degree of safety equal to that which would exist if the tank or reservoir were isolated from the (b) Unless other means are provided, engines.

torque limiting means must be provided for accessory drives located on any component of (c) There must be at least 13 mm (½ in) of the transmission and rotor drive system to clear airspace between each tank and each prevent damage to these components from firewall or shroud isolating that tank, unless excessive accessory load. equivalent means are used to prevent heat transfer from each engine compartment to the flammable fluid.

POWERPLANT FIRE PROTECTION (d) Absorbent materials close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids.

CS 27.1183 Lines, fittings, and components (a) Except as provided in sub paragraph (b), each line, fitting, and other component CS 27.1187 Ventilation and drainage carrying flammable fluid in any area subject to Each compartment containing any part of the engine fire conditions must be fire resistant, powerplant installation must have provision for except that flammable fluid tanks and supports ventilation and drainage of flammable fluids.

which are part of and attached to the engine The drainage means must be: must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non (a) Effective under conditions expected to fireproof part will not cause leakage or spillage prevail when drainage is needed; and of flammable fluid. Components must be (b) Arranged so that no discharged fluid shielded or located so as to safeguard against will cause an additional fire hazard.

the ignition of leaking flammable fluid. An integral oil sump of less than 24 litres (5.2 Imperial gallons/25 US quart) capacity on a reciprocating engine need not be fireproof nor CS 27.1189 Shut off means be enclosed by a fireproof shield.

(a) There must be means to shut off each (b) Sub paragraph (a) does not apply to: line carrying flammable fluids into the engine compartment, except: (1) Lines, fittings, and components which are already approved as part of a type (1) Lines, fittings, and components certificated engine; and forming an integral part of an engine; (2) Vent and drain lines, and their (2) For oil systems for which all fittings, whose failure will not result in, or components of the system, including oil add to, a fire hazard.

tanks, are fireproof or located in areas not subject to engine fire conditions; and (c) Each flammable fluid drain and vent must discharge clear of the induction system air (3) For reciprocating engine inlet.

installations only, engine oil system lines in installations using engines of less than 8195 cm (500 cubic inches) displacement.

(b) There must be means to guard against CS 27.1185 Flammable fluids inadvertent operation of each shutoff, and to (a) Each fuel tank must be isolated from make it possible for the crew to reopen it in the engines by a firewall or shroud.

flight after it has been closed.

Amendment 2 1–E–17 Annex to ED Decision 2008/009/R CS 27 BOOK 1 (c) Each shut off valve and its control must (e) Each part of the cowling or engine be designed, located, and protected to function compartment covering subject to high properly under any condition likely to result temperatures due to its nearness to exhaust from an engine fire. system parts or exhaust gas impingement must be fireproof.

(f) A means of retaining each openable or readily removable panel, cowling, or engine or CS 27.1191 Firewalls rotor drive system covering must be provided to (a) Each engine, including the combustor, preclude hazardous damage to rotors or critical turbine, and tailpipe sections of turbine engines control components in the event of structural or must be isolated by a firewall, shroud or mechanical failure of the normal retention equivalent means, from personnel means, unless such failure is extremely compartments, structures, controls, rotor improbable.

mechanisms, and other parts that are: (1) Essential to a controlled landing; and CS 27.1194 Other surfaces (2) Not protected under CS 27.861.

All surfaces aft of, and near, powerplant (b) Each auxiliary power unit and compartments, other than tail surfaces not combustion heater, and any other combustion subject to heat, flames, or sparks emanating equipment to be used in flight, must be isolated from a powerplant compartment, must be at from the rest of the rotorcraft by firewalls, least fire resistant.

shrouds, or equivalent means.

(c) In meeting sub paragraphs (a) and (b), account must be taken of the probable path of a CS 27.1195 Fire detector systems fire as affected by the airflow in normal flight and in autorotation.

Each turbine engine powered rotorcraft must have approved quick acting fire detectors in (d) Each firewall and shroud must be numbers and locations insuring prompt constructed so that no hazardous quantity of air, detection of fire in the engine compartment fluids, or flame can pass from any engine which cannot be readily observed in flight by compartment to other parts of the rotorcraft.

the pilot in the cockpit.

(e) Each opening in the firewall or shroud must be sealed with close fitting, fireproof grommets, bushings, or firewall fittings.

(f) Each firewall and shroud must be fireproof and protected against corrosion.

CS 27.1193 Cowling and engine compartment covering (a) Each cowling and engine compartment covering must be constructed and supported so that it can resist the vibration, inertia, and air loads to which it may be subjected in operation.

(b) There must be means for rapid and complete drainage of each part of the cowling or engine compartment in the normal ground and flight attitudes.

(c) No drain may discharge where it might cause a fire hazard.

(d) Each cowling and engine compartment covering must be at least fire resistant.

Amendment 2 1–E–18 Annex to ED Decision 2008/009/R CS–27 BOOK 1 SUBPART F – EQUIPMENT gearboxes essential to rotor phasing) having an oil GENERAL system independent of the engine oil system.

CS 27.1301 Function and installation (g) An oil pressure warning device to indicate when the pressure falls below a safe Each item of installed equipment must: value in each pressure lubricated main rotor (a) Be of a kind and design appropriate to drive gearbox (including any gearboxes essential to rotor phasing) having an oil system its intended function; independent of the engine oil system.

(b) Be labelled as to its identification, (h) An oil pressure indicator for each function, or operating limitations, or any applicable combination of these factors; engine.

(i) An oil quantity indicator for each oil (c) Be installed according to limitations specified for that equipment; and tank.

(d) Function properly when installed. (j) An oil temperature indicator for each engine.

(k) At least one tachometer to indicate the CS 27.1303 Flight and navigation rpm of each engine and, as applicable: instruments (1) The rpm of the single main rotor; The following are the required flight and navigation instruments: (2) The common rpm of any main rotors whose speeds cannot vary appreciably (a) An airspeed indicator.

with respect to each other; or (b) An altimeter.

(3) The rpm of each main rotor whose speed can vary appreciably with (c) A magnetic direction indicator.

respect to that of another main rotor.

(l) A low fuel warning device for each fuel CS 27.1305 Powerplant instruments tank which feeds an engine. This device must: The following are the required powerplant (1) Provide a warning to the flight instruments: crew when approximately 10 minutes of usable fuel remains in the tank; and (a) A carburettor air temperature indicator, for each engine having a pre heater that can (2) Be independent of the normal fuel provide a heat rise in excess of 33°C (60°F).

quantity indicating system.

(b) A cylinder head temperature indicator, (m) Means to indicate to the flight crew the for each: failure of any fuel pump installed to show (1) Air cooled engine; compliance with CS 27.955.

(n) A gas temperature indicator for each (2) Rotorcraft with cooling shutters; and turbine engine.

(o) Means to enable the pilot to determine (3) Rotorcraft for which compliance with CS 27.1043 is shown in any condition the torque of each turboshaft engine, if a torque other than the most critical flight condition limitation is established for that engine under CS 27.1521 (e).

with respect to cooling.

(c) A fuel pressure indicator, for each (p) For each turbine engine, an indicator to indicate the functioning of the powerplant ice pump fed engine.

protection system.

(d) A fuel quantity indicator, for each fuel (q) An indicator for the fuel filter required tank.

by CS 27.997 to indicate the occurrence of (e) A manifold pressure indicator, for each contamination of the filter at the degree altitude engine.

established by the applicant in compliance with CS 27.955.

(f) An oil temperature warning device to indicate when the temperature exceeds a safe value (r) For each turbine engine, a warning in each main rotor drive gearbox (including any means for the oil strainer or filter required by Amendment 2 1–F–1 Annex to ED Decision 2008/009/R CS–27 BOOK 1 CS 27.1019, if it has no by pass, to warn the this CS–27 must be designed and installed to pilot of the occurrence of contamination of the ensure that they perform their intended strainer or filter before it reaches the capacity functions under any foreseeable operating established in accordance with CS 27.1019 condition.

(a)(2).

(b) The equipment, systems, and (s) An indicator to indicate the proper installations of a multi engine rotorcraft must functioning of any selectable or controllable be designed to prevent hazards to the rotorcraft heater used to prevent ice clogging of fuel in the event of a probable malfunction or system components. failure.

(t) For rotorcraft for which a 30s econd/2 (c) The equipment, systems, and minute OEI power rating is requested, a means installations of single engine rotorcraft must be must be provided to alert the pilot when the engine designed to minimise hazards to the rotorcraft is at the 30s econd and 2m inute OEI power levels, in the event of a probable malfunction or when the event begins, and when the time interval failure.

expires.

(d) In showing compliance with sub (u) For each turbine engine utilising 30 paragraph (a), (b), or (c), the effects of second/2 minute OEI power, a device or system lightning strikes on the rotorcraft must be must be provided for use by ground personnel considered in accordance with CS 27.610.

which: (1) Automatically records each usage INSTRUMENTS: INSTALLATION and duration of power in the 30 second and 2 minute OEI levels; (2) Permits retrieval of the recorded CS 27.1321 Arrangement and visibility data; (a) Each flight, navigation, and powerplant (3) Can be reset only by ground instrument for use by any pilot must be easily visible to him.

maintenance personnel: and (4) Has a means to verify proper (b) For each multi engine rotorcraft, identical powerplant instruments must be operation of the system or device.

located so as to prevent confusion as to which (v) Warning or caution devices to signal to engine each instrument relates.

the flight crew when ferromagnetic particles are detected by the chip detector required by (c) Instrument panel vibration may not damage, or impair the readability or accuracy 27.1337(e).

of, any instrument.

[Amdt 27/2] (d) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting CS 27.1307 Miscellaneous equipment conditions.

The following is the required miscellaneous equipment: CS 27.1322 Warning, caution, and (a) An approved seat for each occupant.

advisory lights (b) An approved safety belt for each occupant. If warning, caution or advisory lights are installed in the cockpit, they must, unless (c) A master switch arrangement.

otherwise approved the Agency, be: (d) An adequate source of electrical energy, (a) Red, for warning lights (lights where electrical energy is necessary for indicating a hazard which may require operation of the rotorcraft.

immediate corrective action); (e) Electrical protective devices.

(b) Amber, for caution lights (lights indicating possible need for future corrective action); CS 27.1309 Equipment, systems, and (c) Green, for safe operation lights; and installations (d) Any other colour, including white, for (a) The equipment, systems, and lights not described in sub paragraphs (a) to (c), installations whose functioning is required by Amendment 2 1–F–2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 provided the colour differs sufficiently from the (2) Both sources cannot be blocked colours prescribed in sub paragraphs (a) to (c) off simultaneously.

to avoid possible confusion.

(d) For unpressurised rotorcraft, sub paragraph (c)(1) does not apply if it can be demonstrated that the static pressure system calibration, when either static pressure source is CS 27.1323 Airspeed indicating system selected is not changed by the other static pressure source being open or blocked.

(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 CS 27.1327 Magnetic direction indicator error when the corresponding pitot and static (a) Except as provided in subpa ragraph (b): pressures are applied.

(1) Each magnetic direction indicator (b) The airspeed indicating system must be must be installed so that its accuracy is not calibrated in flight at forward speeds of 37 km/h excessively affected by the rotorcraft’s (20 knots) and over.

vibration or magnetic fields; and (c) At each forward speed above 80% of (2) The compensated installation may the climbout speed, the airspeed indicator must not have a deviation, in level flight, greater indicate true airspeed, at sea level with a than 10° on any heading.

standard atmosphere, to within an allowable installation error of not more than the greater (b) A magnetic non stabilised direction of: indicator may deviate more than 10° due to the operation of electrically powered systems such (1) ±3% of the calibrated airspeed; or as electrically heated windshields if either a (2) 9.3 km/h (5 knots). magnetic stabilised direction indicator, which does not have a deviation in level flight greater than 10° on any heading, or a gyroscopic direction indicator, is installed. Deviations of a CS 27.1325 Static pressure systems magnetic non stabilised direction indicator of (a) Each instrument with static air case more than 10° must be placarded in accordance connections must be vented so that the influence of with CS 27.1547 (e).

rotorcraft speed, the opening and closing of windows, airflow variation, and moisture or other foreign matter does not seriously affect its CS 27.1329 Automatic pilot system accuracy.

(a) Each automatic pilot system must be (b) Each static pressure port must be designed so that the automatic pilot can: designed and located in such a manner that the correlation between air pressure in the static (1) Be sufficiently overpowered by one pilot to allow control of the rotorcraft; pressure system and true ambient atmospheric static pressure is not altered when the rotorcraft and encounters icing conditions. An antii cing means (2) Be readily and positively or an alternate source of static pressure may be disengaged by each pilot to prevent it from used in showing compliance with this requirement.

interfering with control of the rotorcraft.

If the reading of the altimeter, when on the alternate static pressure system, differs from the (b) Unless there is automatic reading of the altimeter when on the primary static synchronisation, each system must have a system by more than 15 m (50 feet), a correction means to readily indicate to the pilot the card must be provided for the alternate static alignment of the actuating device in relation to system. the control system it operates.

(c) Except as provided in sub paragraph (c) Each manually operated control for the (d), if the static pressure system incorporates system’s operation must be readily accessible to both a primary and an alternate static pressure the pilots.

source, the means for selecting one or the other (d) The system must be designed and source must be designed so that: adjusted so that, within the range of adjustment (1) When either source is selected, available to the pilot, it cannot produce the other is blocked off, and hazardous loads on the rotorcraft or create hazardous deviations in the flight path under Amendment 2 1–F–3 Annex to ED Decision 2008/009/R CS–27 BOOK 1 any flight condition appropriate to its use, one fuel quantity indicator must be installed; either during normal operation or in the event and of a malfunction, assuming that corrective (3) Each exposed sight gauge used as action begins within a reasonable period of a fuel quantity indicator must be protected time.

against damage.

(e) If the automatic pilot integrates signals (c) Fuel flow meter system. If a fuel flow from auxiliary controls or furnishes signals for meter system is installed, each metering operation of other equipment, there must be component must have a means for bypassing the positive interlocks and sequencing of fuel supply if malfunction of that component engagement to prevent improper operation.

severely restricts fuel flow.

(f) If the automatic pilot system can be (d) Oil quantity indicator. There must be coupled to airborne navigation equipment, means means to indicate the quantity of oil in each must be provided to indicate to the pilots the tank: current mode of operation. Selector switch position (1) On the ground (including during is not acceptable as a means of indication.

the filling of each tank); and (2) In flight, if there is an oil transfer CS 27.1335 Flight director systems system or reserve oil supply system.

If a flight director system is installed, means (e) Rotor drive system transmissions and must be provided to indicate to the flight crew its gearboxes utilising ferromagnetic materials current mode of operation. Selector switch position must be equipped with chip detectors designed is not acceptable as a means of indication.

to indicate the presence of ferromagnetic particles resulting from damage or excessive wear. Chip detectors must: CS 27.1337 Powerplant instruments (1) be designed to provide a signal to (a) Instruments and instrument lines the indicator required by 27.1305(v); and (1) Each powerplant instrument line (2) be provided with a means to allow must meet the requirements of CS 27.961 and crew members to check, in flight, the 27.993.

function of each detector electrical circuit (2) Each line carrying flammable and signal.

fluids under pressure must: (i) Have restricting orifices or other safety devices at the source of pressure to prevent the escape of ELECTRICAL SYSTEMS AND EQUIPMENT excessive fluid if the line fails; and (ii) Be installed and located so that the escape of fluids would not CS 27.1351 General create a hazard.

(a) Electrical system capacity. Electrical (3) Each powerplant instrument that equipment must be adequate for its intended utilises flammable fluids must be installed use. In addition: and located so that the escape of fluid would not create a hazard. (1) Electric power sources, their transmission cables, and their associated (b) Fuel quantity indicator. Each fuel control and protective devices must be able to quantity indicator must be installed to clearly furnish the required power at the proper indicate to the flight crew the quantity of fuel in voltage to each load circuit essential for safe each tank in flight. In addition: operation; and (1) Each fuel quantity indicator must (2) Compliance with paragraph (a) be calibrated to read ‘zero’ during level (1) must be shown by an electrical load flight when the quantity of fuel remaining in analysis, or by electrical measurements that the tank is equal to the unusable fuel supply take into account the electrical loads applied determined under CS 27.959; to the electrical system, in probable (2) When two or more tanks are combinations and for probable durations.

closely interconnected by a gravity feed (b) Function. For each electrical system system and vented, and when it is impossible the following apply: to feed from each tank separately, at least Amendment 2 1–F–4 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (1) Each system, when installed, must ensure that no external power supply having a be: reverse polarity, or a reverse phase sequence, can supply power to the rotorcraft’s electrical (i) Free from hazards in itself, system.

in its method of operation, and in its effects on other parts of the rotorcraft; and CS 27.1353 Storage battery design and (ii) Protected from fuel, oil, installation water, other detrimental substances, (a) Each storage battery must be designed and mechanical damage.

and installed as prescribed in this paragraph.

(2) Electric power sources must (b) Safe cell temperatures and pressures function properly when connected in must be maintained during any probable combination or independently.

charging and discharging condition. No (3) No failure or malfunction of any uncontrolled increase in cell temperature may source may impair the ability of any result when the battery is recharged (after remaining source to supply load circuits previous complete discharge): essential for safe operation.

(1) At maximum regulated voltage or (4) Each electric power source power; control must allow the independent operation (2) During a flight of maximum of each source.

duration; and (c) Generating system . There must be at (3) Under the most adverse cooling least one generator if the system supplies power condition likely to occur in service.

to load circuits essential for safe operation. In (c) Compliance with sub paragraph (b) addition: must be shown by test unless experience with (1) Each generator must be able to similar batteries and installations has shown deliver its continuous rated power; that maintaining safe cell temperatures and (2) Generator voltage control pressures presents no problem.

equipment must be able to dependably (d) No explosive or toxic gases emitted by regulate each generator output within rated any battery in normal operation, or as the result limits; of any probable malfunction in the charging (3) Each generator must have a system or battery installation, may accumulate reverse current cut out designed to in hazardous quantities within the rotorcraft.

disconnect the generator from the battery and (e) No corrosive fluids or gases that may from the other generators when enough escape from the battery may damage reverse current exists to damage that surrounding structures or adjacent essential generator; and equipment.

(4) Each generator must have an over (f) Each nickel cadmium battery voltage control designed and installed to installation capable of being used to start an prevent damage to the electrical system, or to engine or auxiliary power unit must have equipment supplied by the electrical system, provisions to prevent any hazardous effect on that could result if that generator were to structure or essential systems that may be develop an over voltage condition.

caused by the maximum amount of heat the battery can generate during a short circuit of the (d) Instruments. There must be means to battery or of its individual cells.

indicate to appropriate crew members the electric power system quantities essential for (g) Nickel cadmium battery installations safe operation of the system. In addition – capable of being used to start an engine or auxiliary power unit must have: (1) For direct current systems, an ammeter that can be switched into each (1) A system to control the charging generator feeder may be used; and rate of the battery automatically so as to prevent battery overheating; (2) If there is only one generator, the ammeter may be in the battery feeder. (2) A battery temperature sensing and over temperature warning system with a (e) External power. If provisions are made means for disconnecting the battery from its for connecting external power to the rotorcraft, charging source in the event of an over and that external power can be electrically temperature condition; or connected to equipment other than that used for engine starting, means must be provided to Amendment 2 1–F–5 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (3) A battery failure sensing and (a) Each electric connecting cable must be warning system with a means for of adequate capacity.

disconnecting the battery from its charging (b) Each cable that would overheat in the source in the event of battery failure.

event of circuit overload or fault must be at least flame resistant and may not emit dangerous quantities of toxic fumes.

CS 27.1357 Circuit protective devices (c) Insulation on electrical wire and cable installed in the rotorcraft must be self (a) Protective devices, such as fuses or extinguishing when tested in accordance with circuit breakers, must be installed in each CS–25, appendix F, part I (a)(3).

electrical circuit other than: (1) The main circuits of starter motors; and CS 27.1367 Switches (2) Circuits in which no hazard is Each switch must be: presented by their omission.

(a) Able to carry its rated current; (b) A protective device for a circuit essential to flight safety may not be used to (b) Accessible to the crew; and protect any other circuit.

(c) Labelled as to operation and the circuit (c) Each resettable circuit protective device controlled.

(‘trip free’ device in which the tripping mechanism cannot be overridden by the operating control) must be designed so that: LIGHTS (1) A manual operation is required to restore service after tripping; and CS 27.1381 Instrument lights (2) If an overload or circuit fault exists, the device will open the circuit The instrument lights must: regardless of the position of the operating control. (a) Make each instrument, switch, and other devices for which they are provided easily (d) If the ability to reset a circuit breaker readable; and or replace a fuse is essential to safety in flight, that circuit breaker or fuse must be located and (b) Be installed so that: identified so that it can be readily reset or (1) Their direct rays are shielded replaced in flight.

from the pilot’s eyes; and (e) If fuses are used, there must be one spare of each rating, or 50% spare fuses of each (2) No objectionable reflections are visible to the pilot.

rating, whichever is greater.

CS 27.1361 Master switch CS 27.1383 Landing lights (a) There must be a master switch (a) Each required landing or hovering light arrangement to allow ready disconnection of must be approved.

each electric power source from the main bus.

(b) Each landing light must be installed so The point of disconnection must be adjacent to that: the sources controlled by the switch.

(b) Load circuits may be connected so that (1) No objectionable glare is visible they remain energised after the switch is to the pilot; opened, if they are protected by circuit (2) The pilot is not adversely affected protective devices, rated at five amperes or less, by halation; and adjacent to the electric power source.

(3) It provides enough light for night (c) The master switch or its controls must operation, including hovering and landing.

be installed so that the switch is easily discernible and accessible to a crew member in (c) At least one separate switch must be flight.

provided, as applicable: (1) For each separately installed landing light; and CS 27.1365 Electric cables Amendment 2 1–F–6 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (2) For each group of landing lights axis, as viewed when looking aft along the installed at a common location. longitudinal axis.

(e) If the rear position light, when mounted as far aft as practicable in accordance with CS 27.1385 Position light system 27.1385 (c), cannot show unbroken light within installation dihedral angle A (as defined in sub paragraph (d)), a solid angle or angles of obstructed (a) General . Each part of each position light visibility totalling not more than 0.04 steradians system must meet the applicable requirements of is allowable within that dihedral angle, if such this paragraph, and each system as a whole must solid angle is within a cone whose apex is at the meet the requirements of CS 27.1387 to 27.1397.

rear position light and whose elements make an angle of 30° with a vertical line passing through (b) Forward position lights. Forward position lights must consist of a red and a green the rear position light.

light spaced laterally as far apart as practicable and installed forward on the rotorcraft so that, CS 27.1389 Position light distribution with the rotorcraft in the normal flying position, the red light is on the left side and the green and intensities light is on the right side. Each light must be (a) General . The intensities prescribed in approved.

this paragraph must be provided by new equipment with light covers and colour filters in (c) Rear position light. The rear position place. Intensities must be determined with the light must be a white light mounted as far aft as light source operating at a steady value equal to practicable, and must be approved.

the average luminous output of the source at the (d) Circuit. The two forward position lights normal operating voltage of the rotorcraft. The and the rear position light must make a single light distribution and intensity of each position circuit.

light must meet the requirements of sub paragraph (b).

(e) Light covers and colour filters. Each light cover or colour filter must be at least (b) Forward and rear position lights. The flame resistant and may not change colour or light distribution and intensities of forward and shape or lose any appreciable light transmission rear position lights must be expressed in terms during normal use.

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 CS 27.1387 Position light system meet the following requirements: dihedral angles (1) Intensities in the horizontal (a) Except as provided in subpa ragraph (e), plane. Each intensity in the horizontal plane each forward and rear position light must, as (the plane containing the longitudinal axis of installed, show unbroken light within the dihedral the rotorcraft and perpendicular to the plane angles described in this paragraph.

of symmetry of the rotorcraft) must equal or (b) Dihedral angle L (left) is formed by two exceed the values in CS 27.1391.

intersecting vertical planes, the first parallel to (2) Intensities in any vertical plane.

the longitudinal axis of the rotorcraft, and the Each intensity in any vertical plane (the other at 110° to the left of the first, as viewed plane perpendicular to the horizontal plane) when looking forward along the longitudinal must equal or exceed the appropriate value in axis.

CS 27.1393, where I is the minimum intensity prescribed in CS 27.1391 for the (c) Dihedral angle R (right) is formed by corresponding angles in the horizontal plane.

two intersecting vertical planes, the first parallel to the longitudinal axis of the (3) Intensities in overlaps between rotorcraft, and the other at 110° to the right of adjacent signals. No intensity in any overlap the first, as viewed when looking forward along between adjacent signals may exceed the the longitudinal axis.

values in CS 27.1395, except that higher intensities in overlaps may be used with main (d) Dihedral angle A (aft) is formed by two beam intensities substantially greater than intersecting vertical planes making angles of the minima specified in CS 27.1391 and 70° to the right and to the left, respectively, to a 27.1393, if the overlap intensities in relation vertical plane passing through the longitudinal to the main beam intensities do not adversely affect signal clarity. When the peak intensity Amendment 2 1–F–7 Annex to ED Decision 2008/009/R CS–27 BOOK 1 Area A Area B Overlaps of the forward position lights is greater than (candelas) (candelas) 100 candelas, the maximum overlap Green in dihedral angle L 10 1 intensities between them may exceed the Red in dihedral angle R 10 1 values in CS 27.1395 if the overlap intensity Green in dihedral angle A 5 1 in Area A is not more than 10% of peak Red in dihedral angle A 5 1 position light intensity and the overlap Rear white in dihedral angle L 5 1 intensity in Area B is not more than 2.5% of Rear white in dihedral angle R 5 1 peak position light intensity.

Where: (a) Area A includes all directions in the adjacent dihedral angle that pass through the CS 27.1391 Minimum intensities in the light source and intersect the common boundary horizontal plane of forward plane at more than 10° but less than 20°; and and rear position lights (b) Area B includes all directions in the Each position light intensity must equal or adjacent dihedral angle that pass through the exceed the applicable values in the following light source and intersect the common boundary table: plane at more than 20°.

Angle from CS 27.1397 Colour specifications right or left Intensity Dihedral angle of longitudinal (candelas) Each position light colour must have the (light included) axis, measured applicable International Commission on from dead ahead Illumination chromaticity co ordinates as L and R (forward red 0° to 10° 40 follows: and green) 10° to 20° 30 20° to 110° 5 (a) Aviation red: A (rear white) 110° to 180° 20 ‘y’ is not greater than 0.335; and ‘z’ is not greater than 0.002.

(b) Aviation green: CS 27.1393 Minimum intensities in any ‘x’ is not greater than 0.440–0.320y; vertical plane of forward and rear position lights ‘x’ is not greater than y–0.170; and Each position light intensity must equal or ‘y’ is not less than 0.390–0.170x.

exceed the applicable values in the following (c) Aviation white: table: ‘x’ is not less than 0.300 and not greater than 0.540; Angle above or below the Intensity ‘y’ is not less than ‘x–0.040’ or ‘y –0.010’, o horizontal plane – whichever is the smaller; and 0 1.0 I ‘y’ is not greater than ‘x + 0.020’ nor ‘0.636– 0° to 5° 0.90 I 0.400x’; 5° to 10° 0.80 I 10° to 15° 0.70 I Where ‘y ’ is the ‘y’ co ordinate of the o 15° to 20° 0.50 I Planckian radiator for the value of ‘x’ 20° to 30° 0.30 I considered.

30° to 40° 0.10 I 40° to 90° 0.05 I CS 27.1399 Riding light CS 27.1395 Maximum intensities in (a) Each riding light required for water overlapping beams of operation must be installed so that it can: forward and rear position (1) Show a white light for at least lights 3.7 km (two nautical miles) at night under No position light intensity may exceed the clear atmospheric conditions; and applicable values in the following table, except (2) Show a maximum practicable as provided in CS 27.1389(b)(3): unbroken light with the rotorcraft on the water.

Maximum intensity Amendment 2 1–F–8 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (b) Externally hung lights may be used. t –t = flash time interval (seconds).

2 1 Normally, the maximum value of effective intensity is obtained when t and t are chosen 2 1 so that the effective intensity is equal to the CS 27.1401 Anti collision light system instantaneous intensity at t and t .

2 1 (a) General . If certification for night (f) Minimum effective intensities for anti operation is requested, the rotorcraft must have collision light. Each anti collision light an anti collision light system that: effective intensity must equal or exceed the (1) Consists of one or more approved applicable values in the following table: anti collision lights located so that their emitted light will not impair the crew’s Angle above or below Effective intensity vision or detract from the conspicuity of the the horizontal plane (candelas) position lights; and 0° to 5° 150 (2) Meets the requirements of sub 5° to 10° 90 paragraphs (b) to (f).

10° to 20° 30 20° to 30° 15 (b) Field of coverage. The system must consist of enough lights to illuminate the vital areas around the rotorcraft, considering the SAFETY EQUIPMENT 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 CS 27.1411 General rotorcraft, except that there may be solid angles (a) Required safety equipment to be used of obstructed visibility totalling not more than by the crew in an emergency, such as flares and 0.5 steradians.

automatic liferaft releases, must be readily (c) Flashing characteristics . The accessible.

arrangement of the system, that is, the number (b) Stowage provisions for required safety of light sources, beam width, speed of rotation, equipment must be furnished and must: and other characteristics, must give an effective flash frequency of not less than 40, nor more (1) Be arranged so that the equipment than 100, cycles per minute. The effective flash is directly accessible and its location is frequency is the frequency at which the obvious; and rotorcraft’s complete anti collision light system (2) Protect the safety equipment from is observed from a distance, and applies to each damage caused by being subjected to the sector of light including any overlaps that exist inertia loads specified in CS 27.561.

when the system consists of more than one light t CS 27.1413 Safety belts I(t)dt

Ú

t I = Each safety belt must be equipped with a e ) t (t 2 0 - + ◊ 1 2 metal to metal latching device.

source. In overlaps, flash frequencies may exceed 100, but not 180, cycles per minute.

CS 27.1415 Ditching equipment (d) Colour . Each anti collision light must be aviation red and must meet the applicable (a) Emergency flotation and signalling requirements of CS 27.1397.

equipment required by any applicable operating rule must meet the requirements of this (e) Light intensity . The minimum light paragraph.

intensities in any vertical plane, measured with the red filter (if used) and expressed in terms of (b) Each raft and each life preserver must ‘effective’ intensities, must meet the be approved and must be installed so that it is requirements of sub paragraph (f). The readily available to the crew and passengers.

following relation must be assumed: The storage provisions for life preservers must where: accommodate one life preserver for each occupant for which certification for ditching is I = effective intensity (candelas).

e requested.

I(t) = instantaneous intensity as a function of time.

Amendment 2 1–F–9 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (c) Each raft released automatically or by contain information necessary for safe operation the pilot must be attached to the rotorcraft by a of the rotorcraft in icing conditions.

line to keep it alongside the rotorcraft. This line must be weak enough to break before submerging the empty raft to which it is CS 27.1435 Hydraulic systems attached.

(a) Design . Each hydraulic system and its (d) Each signalling device must be free elements must withstand, without yielding, any from hazard in its operation and must be structural loads expected in addition to installed in an accessible location.

hydraulic loads.

(b) Tests . Each system must be substantiated by proof pressure tests. When CS 27.1419 Ice protection proof tested, no part of any system may fail, (a) To obtain certification for flight into malfunction, or experience a permanent set.

icing conditions, compliance with this The proof load of each system must at least 1.5 paragraph must be shown.

times the maximum operating pressure of that system.

(b) It must be demonstrated that the (c) Accumulators . No hydraulic rotorcraft can be safely operated in the accumulator or pressurised reservoir may be continuous maximum and intermittent installed on the engine side of any firewall maximum icing conditions determined under unless it is an integral part of an engine.

appendix C of CS–29 within the rotorcraft altitude envelope. An analysis must be performed to establish, on the basis of the CS 27.1457 Cockpit voice recorders rotorcraft’s operational needs, the adequacy of the ice protection system for the various (a) Each cockpit voice recorder required by components of the rotorcraft.

the applicable operating rules must be approved, and must be installed so that it will (c) In addition to the analysis and physical record the following: evaluation prescribed in sub paragraph (b), the effectiveness of the ice protection system and its (1) Voice communications components must be shown by flight tests of the transmitted from or received in the rotorcraft rotorcraft or its components in measured by radio.

atmospheric icing conditions and by one or (2) Voice communications of flight more of the following tests as found necessary to determine the adequacy of the ice protection crew members on the flight deck.

system: (3) Voice communications of flight (1) Laboratory dry air or simulated crew members on the flight deck, using the rotorcraft’s interphone system.

icing tests, or a combination of both, of the components or models of the components.

(4) Voice or audio signals identifying navigation or approach aids introduced into a (2) Flight dry air tests of the ice protection system as a whole, or its headset or speaker.

individual components.

(5) Voice communications of flight crew members using the passenger (3) Flight tests of the rotorcraft or its components in measured simulated icing loudspeaker system, if there is such a system, and if the fourth channel is available in conditions.

accordance with the requirements of sub (d) The ice protection provisions of this paragraph (c) (4) (ii).

paragraph are considered to be applicable primarily to the airframe. Powerplant (b) The recording requirements of sub paragraph (a) (2) may be met: installation requirements are contained in Subpart E of this CS–27.

(1) By installing a cockpit mounted area microphone located in the best position (e) A means must be identified or provided for recording voice communications for determining the formation of ice on critical originating at the first and second pilot parts of the rotorcraft. Unless otherwise stations and voice communications of other restricted, the means must be available for crew members on the flight deck when night time as well as daytime operation. The directed to those stations; or rotorcraft flight manual must describe the means of determining ice formation and must Amendment 2 1–F–10 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (2) By installing a continually (2) There is an automatic means to energised or voice activated lip microphone simultaneously stop the recorder and prevent at the first and second pilot stations. The each erasure feature from functioning, within microphone specified in this paragraph must 10 minutes after crash impact; and be so located and if necessary, the (3) There is an aural or visual means preamplifiers and filters of the recorder must for pre flight checking of the recorder for be adjusted or supplemented so that the proper operation.

recorded communications are intelligible when recorded under flight cockpit noise (e) The record container must be located conditions and played back. The level of and mounted to minimise the probability of intelligibility must be approved by the rupture of the container as a result of crash Agency. Repeated aural or visual playback impact and consequent heat damage to the of the record may be used in evaluating record from fire.

intelligibility.

(f) If the cockpit voice recorder has a bulk (c) Each cockpit voice recorder must be erasure device, the installation must be installed so that the part of the communication designed to minimise the probability of or audio signals specified in sub paragraph (a) inadvertent operation and actuation of the obtained from each of the following sources is device during crash impact.

recorded on a separate channel: (g) Each recorder container must be either (1) For the first channel, from each bright orange or bright yellow.

microphone, headset, or speaker used at the first pilot station.

CS 27.1459 Flight recorders (2) For the second channel, from each microphone, headset, or speaker used at the (a) Each flight recorder required by the second pilot station.

applicable operating rules must be installed so that: (3) For the third channel, from the cockpit mounted area microphone, or the (1) It is supplied with airspeed, continually energised or voice activated lip altitude, and directional data obtained from microphone at the first and second pilot sources that meet the accuracy requirements stations.

of CS 27.1323, CS 27.1325, and 27.1327, as applicable; (4) For the fourth channel, from: (2) The vertical acceleration sensor is (i) Each microphone, headset, rigidly attached, and located longitudinally or speaker used at the stations for the within the approved centre of gravity limits third and fourth crew members; or of the rotorcraft; (ii) If the stations specified in (3) It receives its electrical power sub paragraph (c) (4) (i) are not from the bus that provides the maximum required or if the signal at such a reliability for operation of the flight recorder station is picked up by another channel, without jeopardising service to essential or each microphone on the flight deck that emergency loads; is used with the passenger loud speaker system if its signals are not picked up (4) There is an aural or visual means by another channel.

for pre flight checking of the recorder for proper recording of data in the storage (iii) Each microphone on the medium; flight deck that is used with the rotorcraft’s loudspeaker system if its (5) Except for recorders powered signals are not picked up by another solely by the engine driven electrical channel.

generator system, there is an automatic means to simultaneously stop a recorder that (d) Each cockpit voice recorder must be has a data erasure feature and prevent each installed so that: erasure feature from functioning, within 10 (1) It receives its electric power from minutes after any crash impact; and the bus that provides the maximum reliability (b) Each non ejectable recorder container for operation of the cockpit voice recorder must be located and mounted so as to minimise without jeopardising service to essential or the probability of container rupture resulting emergency loads; Amendment 2 1–F–11 Annex to ED Decision 2008/009/R CS–27 BOOK 1 from crash impact and subsequent damage to the record from fire.

(c) A correlation must be established between the flight recorder readings of airspeed, altitude, and heading and the corresponding readings (taking into account correction factors) of the first pilot’s instruments. This correlation must cover the airspeed range over which the aircraft is to be operated, the range of altitude to which the aircraft is limited, and 360° of heading. Correlation may be established on the ground as appropriate.

(d) Each recorder container must: (1) Be either bright orange or bright yellow; (2) Have a reflective tape affixed to its external surface to facilitate its location underwater; and (3) Have an underwater locating device, when required by the applicable operating rules, on or adjacent to the container which is secured in such a manner that they are not likely to be separated during crash impact.

CS 27.1461 Equipment containing high energy rotors (a) Equipment containing high energy rotors must meet sub paragraphs (b), (c), or (d).

(b) High energy rotors contained in equipment must be able to withstand damage caused by malfunctions, vibration, abnormal speeds, and abnormal temperatures. In addition: (1) Auxiliary rotor cases must be able to contain damage caused by the failure of high energy rotor blades; and INTENTIONALLY LEFT BLANK (2) Equipment control devices, systems, and instrumentation must reasonably ensure that no operating limitations affecting the integrity of high energy rotors will be exceeded in service.

(c) It must be shown by test that equipment containing high energy rotors can contain any failure of a high energy rotor that occurs at the highest speed obtainable with the normal speed control devices inoperative.

(d) Equipment containing high energy rotors must be located where rotor failure will neither endanger the occupants nor adversely affect continued safe flight.

Amendment 2 1–F–12 Annex to ED Decision 2008/009/R CS–27 BOOK 1 SUBPART G – OPERATING LIMITATIONS AND INFORMATION integrating more than one of these variables) GENERAL are used at one time; and (2) The ranges of these variables (or of the indications on instruments integrating more than one of these variables) are large CS 27.1501 General enough to allow an operationally practical and (a) Each operating limitation specified in safe variation of V .

NE CS 27.1503 to 27.1525 and other limitations and (c) For helicopters, a stabilised power off information necessary for safe operation must be V denoted as V (power off) may be NE NE established.

established at a speed less than V established NE (b) The operating limitations and other pursuant to sub paragraph (a), if the following information necessary for safe operation must be conditions are met: made available to the crew members as (1) V (power off) is not less than a NE prescribed in CS 27.1541 to 27.1589.

speed midway between the power on V and NE the speed used in meeting the requirements of: OPERATING LIMITATIONS (i) CS 27.65(b) for single engine helicopters; and (ii) CS 27.67 for multi engine helicopters.

CS 27.1503 Airspeed limitations: general (2) V (power off) is: NE (a) An operating speed range must be established.

(i) A constant airspeed; (b) When airspeed limitations are a function (ii) A constant amount less than of weight, weight distribution, altitude, rotor power on V ; or NE speed, power, or other factors, airspeed (iii) A constant airspeed for a limitations corresponding with the critical portion of the altitude range for which combinations of these factors must be certification is requested, and a constant established.

amount less than power on V for the NE remainder of the altitude range.

CS 27.1505 Never exceed speed (a) The never exceed speed, V , must be CS 27.1509 Rotor speed NE established so that it is: (a) Maximum power off (autorotation). The (1) Not less than 74 km/h (40 knots) maximum power off rotor speed must be (CAS); and established so that it does not exceed 95% of the lesser of: (2) Not more than the lesser of: (1) The maximum design rpm (i) 0.9 times the maximum determined under CS 27.309(b); and forward speeds established under CS 27.309; (2) The maximum rpm shown during the type tests.

(ii) 0.9 times the maximum speed shown under CS 27.251 and 27.629; or (b) Minimum power off. The minimum power off rotor speed must be established so that (iii) 0.9 times the maximum speed it is not less than 105% of the greater of: substantiated for advancing blade tip mach number effects.

(1) The minimum shown during the type tests; and (b) V may vary with altitude, rpm, NE temperature, and weight, if: (2) The minimum determined by design substantiation.

(1) No more than two of these variables (or no more than two instruments 1–G–1 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (c) Minimum power on. The minimum (1) The maximum rotational speed power on rotor speed must be established so that which may not be greater than: it is: (i) The maximum value (1) Not less than the greater of: determined by the rotor design; or (i) The minimum shown during (ii) The maximum value shown the type tests; and during the type tests; (ii) The minimum determined by (2) The minimum rotational speed design substantiation; and shown under the rotor speed requirements in CS 27.1509(c); and (2) Not more than a value determined under CS 27.33 (a)(1) and (b)(l). (3) The gas temperature limits for turbine engines over the range of operating and atmospheric conditions for which certification is requested.

CS 27.1519 Weight and centre of gravity (d) Fuel grade or designation. The minimum fuel grade (for reciprocating engines), The weight and centre of gravity limitations or fuel designation (for turbine engines), must be determined under CS 27.25 and 27.27, established so that it is not less than that required respectively, must be established as operating for operation of the engines within the limitations.

limitations in sub paragraphs (b) and (c).

(e) Turboshaft engine torque. For rotorcraft with main rotors driven by turboshaft engines, and that do not have a torque limiting device in CS 27.1521 Powerplant limitations the transmission system, the following apply: (a) General. The powerplant limitations (1) A limit engine torque must be prescribed in this paragraph must be established established if the maximum torque that the so that they do not exceed the corresponding engine can exert is greater than: limits for which the engines are type certificated.

(i) The torque that the rotor (b) Take off operation. The powerplant drive system is designed to transmit; or take off operation must be limited by: (ii) The torque that the main (1) The maximum rotational speed, rotor assembly is designed to withstand which may not be greater than: in showing compliance with CS 27.547(e).

(i) The maximum value determined by the rotor design; or (2) The limit engine torque established under sub paragraph (e)(1) may not exceed (ii) The maximum value shown either torque specified in sub paragraph during the type tests; (e)(1)(i) or (ii) .

(2) The maximum allowable manifold (f) Ambient temperature. For turbine pressure (for reciprocating engines); engines, ambient temperature limitations (3) The time limit for the use of the (including limitations for winterization power corresponding to the limitations installations, if applicable) must be established as established in sub paragraphs (b)(1) and (2); the maximum ambient atmospheric temperature at which compliance with the cooling provisions (4) If the time limit in sub paragraph of CS 27.1041 to 27.1045 is shown.

(b)(3) exceeds 2 minutes, the maximum (g) Two and one half minute OEI power allowable cylinder head, coolant outlet, or oil operation . Unless otherwise authorised, the use temperatures; of 2½ minute OEI power must be limited to (5) The gas temperature limits for engine failure operation of multi engine, turbine turbine engines over the range of operating powered rotorcraft for not longer that 2½ minutes and atmospheric conditions for which after failure of an engine. The use of 2½ minute certification is requested.

OEI power must also be limited by: (c) Continuous operation. The continuous (1) The maximum rotational speed, operation must be limited by: which may not be greater than: 1–G–2 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (i) The maximum value to not more than 30 seconds for any period in determined by the rotor design; or which that power is used, and by: (ii) The maximum demonstrated (1) The maximum rotational speed during the type tests; which may not be greater than: (2) The maximum allowable gas (i) The maximum value temperature; and determined by the rotor design: or (3) The maximum allowable torque.

(ii) The maximum value demonstrated during the type tests: (h) Thirty minute OEI power operation.

Unless otherwise authorised, the use of 30 (2) The maximum allowable gas minute OEI power must be limited to multi temperature; and engine, turbine powered rotorcraft for not longer (3) The maximum allowable torque.

than 30 minutes after failure of an engine. The use of 30 minute OEI power must also be limited (k) Rated 2 minute OEI power operation.

by: Rated 2 minute OEI power is permitted only on multi engine, turbine powered rotorcraft, also (1) The maximum rotational speed certificated for the use of rated 30 second OEI which may not be greater than: power, and can only be used for continued (i) The maximum value operation of the remaining engine(s) after a determined by the rotor design; or failure or precautionary shutdown of an engine.

It must be shown that following application of 2 (ii) The maximum value minute OEI power, any damage will be readily demonstrated during the type tests; detectable by the applicable inspections and other (2) The maximum allowable gas related procedures furnished in accordance with temperature; and A27.4 of appendix A of this CS–27. The use of 2 minute OEI power must be limited to not more (3) The maximum allowable torque.

than 2 minutes for any period in which that (i) Continuous OEI power operation.

power is used, and by: Unless otherwise authorised, the use of (1) The maximum rotational speed, continuous OEI power must be limited to multi which may not be greater than: engine, turbine powered rotorcraft for continued flight after failure of an engine. The use of (i) The maximum value continuous OEI power must also be limited by: determined by the rotor design; or (1) The maximum rotational speed, (ii) The maximum value which may not be greater than: demonstrated during the type tests; (i) The maximum value (2) The maximum allowable gas determined by the rotor design; or temperature; and (ii) The maximum value (3) The maximum allowable torque.

demonstrated during the type tests; (2) The maximum allowable gas temperature; and CS 27.1523 Minimum flight crew (3) The maximum allowable torque.

(j) Rated 30 second OEI power operation.

The minimum flight crew must be established Rated 30 second OEI power is permitted only on so that it is sufficient for safe operation, multi engine, turbine powered rotorcraft, also considering: certificated for the use of rated 2 minute OEI (a) The workload on individual crew power, and can only be used for continued members; operation of the remaining engine(s) after a failure or precautionary shutdown of an engine.

(b) The accessibility and ease of operation of It must be shown that following application of necessary controls by the appropriate crew 30 second OEI power, any damage will be readily member; and detectable by the applicable inspections and other (c) The kinds of operation authorised under related procedures furnished in accordance with CS 27.1525.

paragraph A27.4 of Appendix A of this CS 27.

The use of 30 second OEI power must be limited 1–G–3 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 CS 27.1525 Kinds of operations (b) Each arc and line must be wide enough, and located, to be clearly visible to the pilot.

The kinds of operations (such as VFR, IFR, day, night, or icing) for which the rotorcraft is approved are established by demonstrated compliance with the applicable certification CS 27.1545 Airspeed indicator requirements and by the installed equipment.

(a) Each airspeed indicator must be marked as specified in sub paragraph (b), with the marks located at the corresponding indicated airspeeds.

CS 27.1527 Maximum operating altitude (b) The following markings must be made: The maximum altitude up to which operation (1) A red radial line: is allowed, as limited by flight, structural, powerplant, functional, or equipment (i) For rotorcraft other than characteristics, must be established. helicopters, at V ; and NE (ii) For helicopters at V NE (power on).

(2) A red cross hatched radial line at CS 27.1529 Instructions for Continued V (power off) for helicopters, if V (power NE NE Airworthiness off) is less than V (power on).

NE Instructions for Continued Airworthiness in (3) For the caution range, a yellow arc.

accordance with Appendix A must be prepared.

(4) For the safe operating range, a green arc.

MARKINGS AND PLACARDS CS 27.1547 Magnetic direction indicator CS 27.1541 General (a) A placard meeting the requirements of (a) The rotorcraft must contain: this paragraph must be installed on or near the magnetic direction indicator.

(1) The markings and placards specified in CS 27.1545 to 27.1565, and (b) The placard must show the calibration of the instrument in level flight with the engines (2) Any additional information, operating.

instrument markings, and placards required for the safe operation of rotorcraft with (c) The placard must state whether the unusual design, operating or handling calibration was made with radio receivers on or characteristics.

off.

(b) Each marking and placard prescribed in (d) Each calibration reading must be in sub paragraph (a): terms of magnetic heading in not more than 45° increments.

(1) Must be displayed in a conspicuous place; and (e) If a magnetic non stabilised direction indicator can have a deviation of more than 10° (2) May not be easily erased, caused by the operation of electrical equipment, disfigured, or obscured.

the placard must state which electrical loads, or combination of loads, would cause a deviation of more than 10° when turned on.

CS 27.1543 Instrument markings: general For each instrument: CS 27.1549 Powerplant instruments (a) When markings are on the cover glass of the instrument, there must be means to maintain For each required powerplant instrument, as the correct alignment of the glass cover with the appropriate to the type of instrument: face of the dial; and 1–G–4 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (a) Each maximum and, if applicable, (3) Each valve control for any engine minimum safe operating limit must be marked of a multi engine rotorcraft must be marked to with a red radial or a red line; indicate the position corresponding to each engine controlled.

(b) Each normal operating range must be marked with a green arc or green line, not (c) Usable fuel capacity must be marked as extending beyond the maximum and minimum follows: safe limits; (1) For fuel systems having no selector (c) Each take off and precautionary range controls, the usable fuel capacity of the system must be marked with a yellow arc or yellow line; must be indicated at the fuel quantity indicator.

(d) Each engine or propeller range that is restricted because of excessive vibration stresses (2) For fuel systems having selector must be marked with red arcs or red lines; and controls, the usable fuel capacity available at each selector control position must be (e) Each OEI limit or approved operating indicated near the selector control.

range must be marked to be clearly differentiated from the markings of sub paragraphs (a) to (d) (d) For accessory, auxiliary, and emergency except that no marking is normally required for controls: the 30 second OEI limit.

(1) Each essential visual position indicator, such as those showing rotor pitch or landing gear position, must be marked so that each crew member can determine at any time CS 27.1551 Oil quantity indicator the position of the unit to which it relates; and Each oil quantity indicator must be marked (2) Each emergency control must be with enough increments to indicate readily and red and must be marked as to method of accurately the quantity of oil.

operation.

(e) For rotorcraft incorporating retractable landing gear, the maximum landing gear operating speed must be displayed in clear view CS 27.1553 Fuel quantity indicator of the pilot.

If the unusable fuel supply for any tank exceeds 3.8 litres (0.8 Imperial gallon/1 US gallon), or 5 % of the tank capacity, whichever is greater, a red arc must be marked on its indicator CS 27.1557 Miscellaneous markings and extending from the calibrated zero reading to the placards lowest reading obtainable in level flight.

(a) Baggage and cargo compartments, and ballast location. Each baggage and cargo compartment and each ballast location must have a placard stating any limitations on contents, CS 27.1555 Control markings including weight, that are necessary under the (a) Each cockpit control, other than primary loading requirements.

flight controls or control whose function is (b) Seats . If the maximum allowable weight obvious, must be plainly marked as to its function to be carried in a seat is less than 77 kg (170 lbs), and method of operation.

a placard stating the lesser weight must be (b) For powerplant fuel controls: permanently attached to the seat structure.

(1) Each fuel tank selector control (c) Fuel and oil filler openings. The must be marked to indicate the position following apply: corresponding to each tank and to each (1) Fuel filler openings must be existing cross feed position; marked at or near the filler cover with: (2) If safe operation requires the use of (i) The word ‘fuel’; any tanks in a specific sequence, that sequence must be marked on, or adjacent to, the selector (ii) For reciprocating engine for those tanks; and powered rotorcraft, the minimum fuel grade; 1–G–5 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (iii) For turbine engine powered (1) Information required by CS rotorcraft, the permissible fuel 27.1583 to 27.1589.

designations; and (2) Other information that is necessary (iv) For pressure fuelling systems, for safe operation because of design, the maximum permissible fuelling operating, or handling characteristics.

supply pressure and the maximum (b) Approved information. Each part of the permissible defuelling pressure.

manual listed in CS 27.1583 to 27.1589, that is (2) Oil filler openings must be marked appropriate to the rotorcraft, must be furnished, at or near the filler cover with the word ‘oil’. verified, and approved, and must be segregated, identified, and clearly distinguished from each (d) Emergency exit placards . Each placard unapproved part of that manual.

and operating control for each emergency exit must be red. A placard must be near each (c) (Reserved).

emergency exit control and must clearly indicate (d) Table of contents. Each rotorcraft flight the location of that exit and its method of manual must include a table of contents if the operation.

complexity of the manual indicates a need for it.

CS 27.1559 Limitations placard CS 27.1583 Operating limitations There must be a placard in clear view of the (a) Airspeed and rotor limitations.

pilot that specifies the kinds of operations (such Information necessary for the marking of as VFR, IFR, day, night or icing) for which the airspeed and rotor limitations on, or near, their rotorcraft is approved.

respective indicators must be furnished. The significance of each limitation and of the colour coding must be explained.

(b) Powerplant limitations. The following CS 27.1561 Safety equipment information must be furnished: (a) Each safety equipment control to be (1) Limitations required by CS operated by the crew in emergency, such as 27.1521.

controls for automatic liferaft releases, must be plainly marked as to its method of operation.

(2) Explanation of the limitations, when appropriate.

(b) Each location, such as a locker or compartment, that carries any fire extinguishing, (3) Information necessary for marking signalling, or other life saving equipment, must the instruments required by CS 27.1549 to be so marked.

27.1553.

(c) Weight and loading distribution. The weight and centre of gravity limits required by CS 27.25 and 27.27, respectively, must be CS 27.1565 Tail rotor furnished. If the variety of possible loading Each tail rotor must be marked so that its disc conditions warrants, instructions must be is conspicuous under normal daylight ground included to allow ready observance of the conditions.

limitations.

(d) Flight crew. When a flight crew of more than one is required, the number and functions of ROTORCRAFT FLIGHT MANUAL AND the minimum flight crew determined under CS APPROVED MANUAL MATERIAL 27.1523 must be furnished.

(e) Kinds of operation. Each kind of operation for which the rotorcraft and its equipment installations are approved must be CS 27.1581 General listed.

(a) Furnishing information . A rotorcraft (f) (Reserved) flight manual must be furnished with each rotorcraft, and it must contain the following: 1–G–6 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (g) Altitude . The altitude established under CS 27.1587 Performance information CS 27.1527 and an explanation of the limiting (a) The Rotorcraft Flight Manual must factors must be furnished.

contain the following information, determined in accordance with CS 27.49 through CS 27.79 and CS 27.143 (c) and (d): CS 27.1585 Operating procedures (1) Enough information to determine the limiting height speed envelope.

(a) Parts of the manual containing operating (2) Information relative to: procedures must have information concerning any normal and emergency procedures and other (i) The steady rates of climb and information necessary for safe operation, descent, in ground effect and out of including take off and landing procedures and ground effect hovering ceilings, together associated airspeeds. The manual must contain with the corresponding airspeeds and any pertinent information including: other pertinent information including the calculated effects of altitude and (1) The kind of take off surface used in temperatures; the tests and each appropriate climb out speed; and (ii) The maximum weight for each altitude and temperature condition (2) The kind of landing surface used in at which the rotorcraft can safely hover the tests and appropriate approach and glide airspeeds. in ground effect and out of ground effect in winds of not less than 31 km/h (17 (b) For multi engine rotorcraft, information knots) from all azimuths. This data must identifying each operating condition in which the be clearly referenced to the appropriate fuel system independence prescribed in CS hover charts. In addition, if there are 27.953 is necessary for safety must be furnished, other combinations of weight, altitude together with instructions for placing the fuel and temperature for which performance system in a configuration used to show information is provided and at which the compliance with that paragraph.

rotorcraft cannot land and take off safely with the maximum wind value, those (c) For helicopters for which a V (power NE portions of the operating envelope and off) is established under CS 27.1505 (c), the appropriate safe wind conditions information must be furnished to explain the V NE must be stated in the Rotorcraft Flight (power off) and the procedures for reducing Manual; airspeed to not more than the V (power off) NE following failure of all engines.

(iii) For reciprocating engine powered rotorcraft, the maximum (d) For each rotorcraft showing compliance atmospheric temperature at which with CS 27.1353(g)(2) or (g)(3), the operating compliance with the cooling provisions procedures for disconnecting the battery from its of CS 27.1041 to 27.1045 is shown; and charging source must be furnished.

(iv) Glide distance as a function (e) If the unusable fuel supply in any tank of altitude when autorotating at the exceeds 5 % of the tank capacity, or 3.8 litres speeds and conditions for minimum rate (0.8 Imperial gallon/1 US gallon), whichever is of descent and best glide as determined greater, information must be furnished which in CS 27.71.

indicates that when the fuel quantity indicator reads ‘zero’ in level flight, any fuel remaining in (b) The rotorcraft flight manual must the fuel tank cannot be used safely in flight.

contain: (f) Information on the total quantity of (1) In its performance information usable fuel for each fuel tank must be furnished.

section any pertinent information concerning the take off weights and altitudes used in (g) The airspeeds and rotor speeds for compliance with CS 27.51; and minimum rate of descent and best glide angle as prescribed in CS 27.71 must be provided.

(2) The horizontal take off distance determined in accordance with CS 27.65(a)(2)(i).

[Amdt. No.: 27/1] 1–G–7 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 27.1589 Loading information There must be loading instructions for each possible loading condition between the maximum and minimum weights determined under CS 27.25 that can result in a centre of gravity beyond any extreme prescribed in CS 27.27, assuming any probable occupant weights.

INTENTIONALLY LEFT BLANK 1–G–8 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 APPENDICES Appendix A – Instructions for Continued Airworthiness A27.1 General components and systems are controlled and how they operate, including any special (a) This appendix specifies requirements procedures and limitations that apply.

for the preparation of instructions for continued (4) Servicing information that covers airworthiness as required by CS 27.1529.

details regarding servicing points, capacities (b) The instructions for continued of tanks, reservoirs, types of fluids to be airworthiness for each rotorcraft must include used, pressures applicable to the various the instructions for continued airworthiness for systems, location of access panels for each engine and rotor (hereinafter designated inspection and servicing, locations of ‘products’), for each appliance required by any lubrication points, the lubricants to be used, applicable CS or operating rule, and any equipment required for servicing, tow required information relating to the interface of instructions and limitations, mooring, jacking, and levelling information.

those appliances and products with the rotorcraft. If instructions for continued (b) Maintenance instructions airworthiness are not supplied by the (1) Scheduling information for each manufacturer of an appliance or product part of the rotorcraft and its engines, installed in the rotorcraft the instructions for auxiliary power units, rotors, accessories, continued airworthiness for the rotorcraft must instruments and equipment that provides the include the information essential to the recommended periods at which they should continued airworthiness of the rotorcraft.

be cleaned, inspected, adjusted, tested, and lubricated, and the degree of inspection, the applicable wear tolerances, and work A27.2 Format recommended at these periods. However, it is allowed to refer to an accessory, instrument, (a) The instructions for continued or equipment manufacturer as the source of airworthiness must be in the form of a manual this information if it is shown that the item or manuals as appropriate for the quantity of has an exceptionally high degree of data to be provided.

complexity requiring specialised maintenance techniques, test equipment, or (b) The format of the manual or manuals expertise. The recommended overhaul must provide for a practical arrangement.

periods and necessary cross references to the Airworthiness Limitations section of the manual must also be included. In addition an A27.3 Content inspection program that includes the The contents of the manual or manuals must frequency and extent of the inspections be prepared in a language acceptable to the necessary to provide for the continued Agency. The instructions for continued airworthiness of the rotorcraft must be included.

airworthiness must contain the following manuals or sections, as appropriate, and (2) Troubleshooting information information: describing probable malfunctions, how to recognise those malfunctions, and the (a) Rotorcraft maintenance manual or remedial action for those malfunctions.

section (3) Information describing the order (1) Introduction information that and method of removing and replacing includes an explanation of the rotorcraft’s products and parts with any necessary features and data to the extent necessary for precautions to be taken.

maintenance or preventive maintenance.

(4) Other general procedural (2) A description of the rotorcraft and instructions including procedures for system its systems and installations including its testing during ground running, symmetry engines, rotors, and appliances.

checks, weighing and determining the centre of gravity, lifting and shoring, and storage (3) Basic control and operation limitations.

information describing how the rotorcraft 1–App A–1 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 (c) Diagrams of structural access plates and information needed to gain access for inspections when access plates are not provided.

(d) Details for the application of special inspection techniques including radiographic and ultrasonic testing where such processes are specified.

(e) Information needed to apply protective treatments to the structure after inspection.

(f) All data relative to structural fasteners such as identification, discard recommendations, and torque values.

(g) A list of special tools needed.

[Amdt 27/2] A27.4 Airworthiness Limitations Section INTENTIONALLY LEFT BLANK The instructions for continued airworthiness must contain a section titled airworthiness limitations, that is segregated and clearly distinguishable from the rest of the document.

This section must set forth each mandatory replacement time, structural inspection interval, and related structural inspection procedure approved under CS 27.571. If the instructions for continued airworthiness consist of multiple documents, the section required by this paragraph must be included in the principal manual. This section must contain a legible statement in a prominent location that reads: ‘the airworthiness limitations section is approved and variations must also be approved.’ 1–App A–2 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 Appendix B Airworthiness Criteria for Helicopter Instrument Flight (2) Cruise. Stability must be shown I. General . A small helicopter may not be throughout the speed range from 0.7 to 1.1 V H type certificated for operation under the or V , whichever is lower, not to exceed ±37 NEI instrument flight rules (IFR) unless it meets the km/h (±20 knots) from trim with: design and installation requirements contained in this appendix.

(i) The helicopter trimmed and power adjusted for level flight at 0.9 V H II. Definitions or 0.9 V , whichever is lower; and NEI (a) V means instrument climb speed, YI (ii) Landing gear retracted (if utilised instead of V for compliance with the Y retractable).

climb requirements for instrument flight.

(3) Slow cruise. Stability must be (b) V means instrument flight never NEI shown throughout the speed range from exceed speed, utilised instead of V for NE 0.9 V to 1.3 V or 37 km/h (20 knots) MINI MINI compliance with maximum limit speed above trim speed, whichever is greater, with: requirements for instrument flight.

(i) The helicopter trimmed and (c) V means instrument flight minimum MINI power adjusted for level flight at speed, utilised in complying with minimum limit 1.1 V ; and MINI speed requirements for instrument flight.

(ii) Landing gear retracted (if III. Trim. It must be possible to trim the retractable).

cyclic, collective, and directional control forces (4) Descent. Stability must be shown to zero at all approved IFR airspeeds, power throughout the speed range 37 km/h (20 knots) settings, and configurations appropriate to the either side of trim with: type.

(i) The helicopter trimmed at IV. Static longitudinal stability 0.8 V or 0.8 V (or 0.8 V for the H NEI LE landing gear extended case), whichever (a) General . The helicopter must possess is lower; positive static longitudinal control force stability at critical combinations of weight and centre of (ii) Power required for 1000 fpm gravity at the conditions specified in paragraphs descent at trim speed; and IV (b) or (c) of this Appendix. The stick force (iii) Landing gear extended and must vary with speed so that any substantial retracted, if applicable.

speed change results in a stick force clearly perceptible to the pilot. For single pilot approval (5) Approach. Stability must be shown the airspeed must return to within 10% of the throughout the speed range from 0.7 times the trim speed when the control force is slowly minimum recommended approach speed to 37 released for each trim condition specified in km/h (20 knots) above the maximum paragraph IV(b) of this Appendix.

recommended approach speed with: (b) For single pilot approval (i) The helicopter trimmed at the recommended approach speed or speeds; (1) Climb. Stability must be shown in climb throughout the speed range 37 km/h (20 (ii) Landing gear extended and knots) either side of trim with: retracted, if applicable; and (i) The helicopter trimmed at (iii) Power required to maintain a V ; YI 3° glide path and power required to maintain the steepest approach gradient (ii) Landing gear retracted (if for which approval is requested.

retractable); and (c) Helicopters approved for a minimum (iii) Power required for limit crew of two pilots must comply with the climb rate (at least 5 m/s (1000 fpm)) at provisions of paragraphs IV(b)(2) and IV(b)(5) of V or maximum continuous power, YI this Appendix.

whichever is less.

1–App B–1 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 Appendix B (Continued) V. Static lateral directional stability VII. Stability augmentation system (SAS) (a) Static directional stability must be (a) If a SAS is used, the reliability of the SAS positive throughout the approved ranges of must be related to the effects of its failure. Any SAS airspeed, power, and vertical speed. In straight failure condition that would prevent continued and steady sideslips up to ±10° from trim, safe flight and landing must be extremely directional control position must increase without improbable. It must be shown that, for any failure discontinuity with the angle of sideslip, except condition of the SAS which is not shown to be for a small range of sideslip angles around trim. extremely improbable: At greater angles up to the maximum sideslip (1) The helicopter is safely controllable angle appropriate to the type, increased when the failure or malfunction occurs at any directional control position must produce speed or altitude within the approved IFR increased angle of sideslip. It must be possible to operating limitations; and maintain balanced flight without exceptional (2) The overall flight characteristics pilot skill or alertness.

of the helicopter allow for prolonged (b) During sideslips up to ±10° from trim instrument flight without undue pilot effort.

throughout the approved ranges of airspeed, Additional unrelated probable failures power, and vertical speed there must be no affecting the control system must be negative dihedral stability perceptible to the pilot considered. In addition: through lateral control motion or force.

Longitudinal cyclic movement with sideslip must (i) The controllability and not be excessive.

manoeuvrability requirements in Subpart B of CS 27 must be met throughout a [Amdt. No.: 27/1] practical flight envelope; (ii) The flight control, trim, and VI. Dynamic stability dynamic stability characteristics must (a) For single pilot approval: not be impaired below a level needed to (1) Any oscillation having a period of allow continued safe flight and landing; less than 5 seconds must damp to ½ amplitude and in not more than one cycle.

(iii) The static longitudinal and (2) Any oscillation having a period of static directional stability requirements 5 seconds or more but less than 10 seconds of Subpart B of CS 27 must be met must damp to ½ amplitude in not more than throughout a practical flight envelope.

two cycles.

(b) The SAS must be designed so that it (3) Any oscillation having a period of cannot create a hazardous deviation in flight path 10 seconds or more but less than 20 seconds or produce hazardous loads on the helicopter must be damped.

during normal operation or in the event of (4) Any oscillation having a period of malfunction or failure, assuming corrective 20 seconds or more may not achieve double action begins within an appropriate period of amplitude in less than 20 seconds.

time. Where multiple systems are installed, (5) Any a periodic response may not subsequent malfunction conditions must be achieve double amplitude in less than considered in sequence unless their occurrence is 6 seconds.

shown to be improbable.

(b) For helicopters approved with a [Amdt. No.: 27/1] minimum crew of two pilots: (1) Any oscillation having a period of VIII. Equipment, systems, and installation.

less than 5 seconds must damp to ½ amplitude The basic equipment and installation must in not more than two cycles.

comply with CS 29.1303, 29.1431 and 29.1433, (2) Any oscillation having a period of with the following exceptions and additions: 5 seconds or more but less than 10 seconds (a) Flight and navigation instruments must be damped.

(1) A magnetic gyros tabilised direction (3) Any oscillation having a period of indicator instead of the gyroscopic direction 10 seconds or more may not achieve double indicator required by CS 29.1303 (h); and amplitude in less than 10 seconds.

(2) A standby attitude indicator which meets the requirements of CS 29.1303(g)(1) to 1–App B–2 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 Appendix B (Continued) (7), instead of a rate of turn indicator required failure or combination of failures that is by CS 29.1303(g). For two pilot not shown to be extremely improbable; configurations, one pilot’s primary indicator and may be designated for this purpose. If standby (iv) For singlepi lot configurations, batteries are provided they may be charged instruments which require a static source from the aircraft electrical system if adequate must be provided with a means of isolation is incorporated.

selecting an alternate source and that (b) Miscellaneous requirements source must be calibrated.

(1) Instrument systems and other IX. Rotorcraft flight manual. A rotorcraft systems essential for IFR flight that could be flight manual or rotorcraft flight manual IFR adversely affected by icing must be adequately supplement must be provided and must contain: protected when exposed to the continuous and (a) Limitations . The approved IFR flight intermittent maximum icing conditions envelope, the IFR flight crew composition, the defined in appendix C of CS–29, whether or revised kinds of operation, and the steepest IFR not the rotorcraft is certificated for operation precision approach gradient for which the in icing conditions.

helicopter is approved; (2) There must be means in the generating system to automatically de energise (b) Procedures . Required information for and disconnect from the main bus any power proper operation of IFR systems and the source developing hazardous overvoltage. recommended procedures in the event of stability augmentation or electrical system failures; and (3) Each required flight instrument using a power supply (electric, vacuum, etc.)

(c) Performance . If V differs from V , YI Y must have a visual means integral with the climb performance at V and with maximum YI instrument to indicate the adequacy of the continuous power throughout the ranges of power being supplied.

weight, altitude, and temperature for which approval is requested.

(4) When multiple systems performing like functions are required, each system must be grouped, routed, and spaced so that physical separation between systems is provided to ensure that a single malfunction will not adversely affect more than one system.

(5) For systems that operate the required flight instruments at each pilot’s station: (i) Only the required flight instruments for the first pilot may be connected to that operating system; (ii) Additional instruments, systems, or equipment may not be connected to an operating system for a second pilot unless provisions are made to ensure the continued normal functioning of the required instruments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable; (iii) The equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the instruments will remain available to a pilot, without additional crewmember action, after any single 1–App B–3 Amendment 2 Annex to ED Decision 2008/009/R CS–27 BOOK 1 Appendix C Criteria for Category A C27.1 General. A small multi engine rotorcraft 29.908(a) – Cooling fans.

may not be type certificated for category A operation 29.917(b) – Rotor drive system: Design.

unless it meets the design installation and and (c)(1) performance requirements contained in this appendix in addition to the requirements of this 29.927(c)(1) – Additional tests.

CS 27.

29.953(a) – Fuel system independence.

C27.2 Applicable CS–29 paragraphs . The 29.1027(a) – Transmission and gearboxes: following paragraphs of CS–29 must be met in General.

addition to the requirements of this code: 29.1045(a)(1), – Climb cooling test procedures.

29.45(a) – General.

(b), (c), (d) and (f) and (b)(2) 29.1047(a) – Take off cooling test procedures.

29.49(a) – Performance at minimum operating speed.

29.1181(a) – Designated fire zones: Regions included.

29.51 – Take off data: General.

29.1187(e) – Drainage and ventilation of fire 29.53 – Take off: Category A.

zones.

29.55 – Take off decision point: 29.1189(c) – Shutoff means.

Category A.

29.1191(a)(l) – Firewalls.

29.59 – Take off path: Category A.

29.1193(e) – Cowling and engine compartment 29.60 – Elevated heliport take off path: covering.

Category A.

29.1195(a) – Fire extinguishing systems (one 29.61 – Take off distance: Category A.

and (d) shot).

29.62 – Rejected take off: Category A.

29.1197 – Fire extinguishing agents.

29.64 – Climb: General.

29.1199 – Extinguishing agent containers.

29.65(a) – Climb: AEO.

29.1201 – Fire extinguishing system 29.67(a) – Climb: OEI.

materials.

29.75 – Landing: General.

29.1305(a)(6) – Powerplant instruments.

and (b) 29.77 – Landing decision point: Category A.

29.1309(b)(2)(i) – Equipment, systems and and (d) installations.

29.79 – Landing: Category A.

29.1323(c)(1) – Airspeed indicating system.

29.81 – Landing distance (ground level sites): Category A.

29.1331(b) – Instruments using a power supply.

29.85 – Balked landing: Category A.

29.1351(d)(2) – Additional requirements for Category A rotorcraft (Operation 29.87(a) – Height velocity envelope.

with the normal electrical power 29.547(a) – Main and tail rotor structure.

generating system inoperative.)

and (b) 29.1587(a) – Performance information.

(29.571 – Fatigue evaluation of structure.)

AC Material only: AC 29 2C (See AC 29 2C Change 2 dated 25 April 2006 and Change 2 dated 25 April 2006, AMC material to CS–29) Paragraph AC29.571A.b(2).

29.861(a) – Fire protection of structure, [Amdt 27/2] controls and other parts.

29.901(c) – Powerplant: Installation.

29.903(b), – Engines.

(c) and (e) Amendment 2 1–App C–1

Annex to ED Decision 2008/009/R CS 27 BOOK 2

EASA Certification Specifications

for

SMALL ROTORCRAFT

CS 27

Book 2

Acceptable Means of Compliance

Amendment 2 1 0 1 Annex to ED Decision 2008/009/R CS–27 BOOK 2 AMC 27 General 1. The AMC to CS–27 consists of FAA AC 271 B Change 2 dated 25 April 2006 with the changes/additions given in this Book 2 of CS–27.

2. The primary reference for each of these AMCs is the CS–27 paragraph. Where there is an appropriate paragraph in FAA AC 271 B Change 2 dated 25 April 2006 this is added as a secondary reference.

[Amdt 27/2] AMC 27.351 Yaw manoeuvre conditions 1. Introduction This AMC provides further guidance and acceptable means of compliance to supplement FAA AC 27 1B Change 2 (AC 27.351. § 27.351 (Amendment 272 6) YAWING CONDITIONS), to meet the Agency's interpretation of CS 27.351. As such it should be used in conjunction with the FAA AC but take precedence over it, where stipulated, in the showing of compliance.

Specifically, this AMC addresses two areas where the FAA AC has been deemed by the Agency as being unclear or at variance to the Agency’s interpretation. These areas are as follows: a. Aerodynamic Loads The certification specification CS 27.351 provides a minimum safety standard for the design of rotorcraft structural components that are subjected in flight to critical loads combinations of antit orque system thrust (e.g. tail rotor), inertia and aerodynamics. A typical example of these structural components is the tailboom.

However, compliance with this standard according to FAA AC 271 B Change 2 may not necessarily be adequate for the design of rotorcraft structural components that are principally subjected in flight to significant aerodynamic loads (e.g. vertical empennage, fins, cowlings and doors).

For these components and their supporting structure, suitable design criteria should be developed by the Applicant and agreed with the Agency.

In lieu of acceptable design criteria developed by the applicant, a suitable combination of sideslip angle and airspeed for the design of rotorcraft components subjected to aerodynamic loads may be obtained from a simulation of the yaw manoeuvre of CS 27.351, starting from the initial directional control input specified in CS 27.351(b)(1) and (c)(1), until the rotorcraft reaches the maximum overswing sideslip angle resulting from its motion around the yaw axis.

b. Interaction of System and Structure Maximum displacement of the directional control, except as limited by pilot effort (CS 27.397(a)), is required for the conditions cited in the certification specification. In the load evaluation credit may be taken for consideration of the effects of control system limiting devices.

However, the probability of failure or malfunction of these system(s) should also be considered and if it is shown not to be extremely improbable then further load conditions with the system in the failed state should be evaluated. This evaluation may include Flight Manual Limitations, if failure of the system is reliably indicated to the crew.

A yaw limiting device is a typical example of a system whose failed condition should be investigated in the assessment of the loads requested by CS 27.351.

Amendment 2 2–1 Annex to ED Decision 2008/009/R CS–27 BOOK 2 An acceptable methodology to investigate the effects of all system failures not shown to be extremely improbable on the loading conditions of CS 27.351 is as follows: i) With the system in the failed state and considering any appropriate reconfiguration and flight limitations, it should be shown that the rotorcraft structure can withstand without failure the loading conditions of CS 27.351, when the manoeuvre is performed in accordance with the provisions of this AMC.

ii) The factor of safety to apply to the above specified loading conditions to comply with CS 27.305 is defined in the figure below.

Qj = (Tj)(Pj) where: Tj = Average flight time spent with a failed limiting system j (in hours) Pj = Probability of occurrence of failure of control limiting system j (per hour) Note: If Pj is greater than 1x10 per flight hour then a 1.5 factor of safety should be applied to all limit load conditions evaluated for the system failure under consideration.

[Amdt 27/2] AMC 27.865 Class D (Human External Cargo) for Operations within Europe 1. Introduction This Additional EASA AMC, used in conjunction with FAA guidance on Human External Cargo (HEC), provides an acceptable means of compliance with CS 27.865 for rotorcraft intended for Class D Rotorcraft/Load Combinations (RLC) for the carriage of Human External Cargo (HEC). For all other RLC classes, reference should be made directly to the adopted FAA AC material.

The addition of this AMC has been necessary due to a difference in operational requirements within the USA and Europe and the absence of dedicated material within the FAA AC.

2. Basic Definition and Intended Use A Class D RLC is one where personnel are at some point in the operation transported external to the rotorcraft, and the operator receives compensation from or on behalf of the person(s) being transported. e.g. Transfer of personnel to/from a ship.

3. Certification Considerations See reference in AMC 27 General Amendment 2 2–2 Annex to ED Decision 2008/009/R CS–27 BOOK 2 Class D HEC was originally envisaged for Part 29/CS 29 rotorcraft only. However, CS2 7 rotorcraft which have been shown to comply with the engine isolation specifications of CS2 7 Appendix C are also eligible.

The rotorcraft must be certified for an OEI/OGE hover performance weight, altitude and temperature envelope. This becomes the maximum envelope that can be used for Class D HEC operations.

4. Compliance Procedures 4.1 The rotorcraft is required to meet the Category A engine isolation specifications of CS2 7 Appendix C, and have One Engine Inoperative/Out of Ground Effect (OEI/OGE) hover performance capability in its approved, jettisonable HEC weight, altitude, and temperature envelope.

(i) In determining OEI hover performance, dynamic engine failures should be considered. Each hover verification test should begin from a stabilized hover at the maximum OEI hover weight, at the requested ing rounde ffect (IGE) or OGE skid or wheel height, and with all engines operating. At this point the critical engine should be failed and the aircraft should remain in a stabilized hover condition without exceeding any rotor limits or engine limits for the operating engine(s). As with all performance testing, engine power should be limited to minimum specification power. Engine failures may be simulated by rapidly moving the throttle to idle provided a ‘needle split’ is obtained between the rotor and engine RPM.

(ii) Normal pilot reaction time should be used following the engine failure to maintain the stabilized hover flight condition. When hovering OGE or IGE at maximum OEI hover weight, an engine failure should not result in an altitude loss of more than 10 percent or four (4) feet, whichever is greater, of the altitude established at the time of engine failure. In either case, sufficient power margin should be available from the operating engine(s) to regain the altitude lost during the dynamic engine failure and to transition to forward flight.

(iii) Consideration should also be given to the time required to recover or manoeuvre the Class D external load and to transition into forward flight. For example to winch up and bring aboard personnel in hoisting operations or manoeuvre clear of power lines for fixed strop/basket operations. The time necessary to perform such actions may exceed the short duration OEI power ratings. For example, for a helicopter with a 30sec/2 min rating structure that sustains an engine failure at a height of 40 feet, the time required to res tabilise in a hover, recover the external load (given the hoist speed limitations), and then transition to forward flight (with minimal altitude loss) would likely exceed 30 seconds and a power reduction into the 2 minute rating would be necessary.

(iv) The Rotorcraft Flight Manual (RFM) should contain information that describes the expected altitude loss, any special recovery techniques, and the time increment used for recovery of the external load when establishing maximum weights and wheel or skid heights. The OEI hover chart should be placed in the performance section of the RFM or RFM supplement. Allowable altitude extrapolation for the hover data should not exceed 2000 feet.

4.2 For helicopters that incorporate engine driven generators, the hoist should remain operational following an engine or generator failure. A hoist should not be powered from a bus that is automatically shed following the loss of an engine or generator. Maximum twoe ngine generator loads should be established so that when one engine or generator fails, the remaining generator can assume the entire rotorcraft electrical load (including the maximum hoist electrical load) without exceeding approved limitations.

4.3 The external load attachment means and the personnel carrying device should be shown to meet the specifications of CS 27.865(a) for the proposed operating envelope.

Amendment 2 2–3 Annex to ED Decision 2008/009/R CS–27 BOOK 2 4.4 The rotorcraft is required to be equipped for, or otherwise allow, direct intercommunication under any operational conditions among crew members and the HEC. For RCL Class D operations, twow ay radios or intercoms should be employed.

[Amdt 27/2] AMC MG4 Full Authority Digital Electronic Controls (FADEC) Note: Certification procedures identified in MG4 refer specifically to the FAA regulatory system. For guidance on EASA procedures, reference should be made to Commission Regulation (EC) No 1702/2003 (as amended) (Part2 1), AMC2 0 (and specifically AMC 201 and 203 ) and to EASA internal working procedures, all of which are available on EASA's web site: http://www.easa.europa.eu/ [Amdt 27/2] Amendment 2 2–4