SUPPLEMENTAL TYPE CERTIFICATE - 10059692 REV. 1

Proprietary document. Copies are not controlled. Confirm revision status through the EASA- 1 nternet/1 ntra net. Page 1 of 1 An agency of the European Union JC/DOC/335 ISSUE 1.4 MAJOR CHANGE or REPAIR APPROVAL SHEET J & C Engineering Services Ltd Tel: +44 (0)1763 252900 Fax: +44 (0)8709 223614 Email: sales@jandceng.co.uk Web: www.jandceng.co.uk TITLE XCAM POD INSTALLATION ON PORT OR STARBOARD WING STRUT AIRCRAFT MAKE/TYPE SERIAL NUMBER REGISTRATION CESSNA 172 - - ASSOCIATED TECHNICAL DOCUMENTS DOCUMENT NUMBER ISSUE DESCRIPTION JC/ER/4584 1.5 XCAM POD INSTALLATION ON PORT OR STARBOARD WING STRUT REASON FOR CHANGE SPECIALISM MEANS OF COMPLIANCE CVE DATE S01 AIRWORTHINESS 0 1 2 3 4 5 6 7 8 9 WER JENKINS 25-06-19 S02 RELIABILITY 0 1 2 3 4 5 6 7 8 9 S03 AERODYNAMICS 0 1 2 3 4 5 6 7 8 9 TJ NEARY 25-06-19 S04 AVIONIC SYSTEMS 0 1 2 3 4 5 6 7 8 9 S05 ELECTRICAL SYSTEMS 0 1 2 3 4 5 6 7 8 9 S06 TECH. PUBLICATIONS 0 1 2 3 4 5 6 7 8 9 WER JENKINS 25-06-19 S07 MECH/FLUID SYSTEMS 0 1 2 3 4 5 6 7 8 9 S08 FURNISHINGS 0 1 2 3 4 5 6 7 8 9 S09 STRUCTURES 0 1 2 3 4 5 6 7 8 9 TJ NEARY 25-06-19 S10 INTERIORS 0 1 2 3 4 5 6 7 8 9 MEANS OF COMPLIANCE CODES 0 – Definition 4 – Lab Rig Test 8 – Simulator Test 1 – Drawing & Description 5 – Ground Test 9 – Equipment Qualification 2 – Calculation & Analysis 6 – Flight Test 3 – System Safety Analysis 7 – Inspection/Survey This Change or Repair is prepared by J & C Engineering Services Ltd Design Organisation Approval Number EASA.21J.037 / AD/1853/05 I hereby declare that this Change or Repair complies with the Airworthiness, Environmental Protection and OSD Requirements applicable to the subject aircraft or equipment and has no unsafe features. AUTHORISED SIGNATORY CHANGE/REPAIR NUMBER WER JENKINS (Head of Design) Signature Date 25-06-19 JC/MOD/M4584 ISSUE 1.5 J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 1 of 30 JC/DOC/325 ISSUE 1.4 TITLE XCAM POD INSTALLATION ON PORT OR STARBOARD WING STRUT AIRCRAFT MAKE/TYPE SERIAL NUMBER REGISTRATION CESSNA 172 - - DOCUMENTS AFFECTED CONTENTS MANUAL AFFECTED AMENDMENT RESPONSIBILITY YES NO J & C OPERATOR AFM X X AMM X CMM X ELA X IPC X MEL X SB X SRM X WBM X X WDM X OTHER: Description 1.0 Certification Basis 2.0 Compliance with Requirements 3.0 Accomplishment Instructions 4.0 Tests 5.0 Manuals 6.0 Weight and Balance 7.0 Electrical Load 8.0 Limitations/Concessions 9.0 Continued Airworthiness 10.0 Documents Required 11.0 COMPILED CHECKED APPROVED AW MASTERS WER JENKINS HEAD OF OFFICE OF AIRWORTHINESS WER JENKINS DATE: 25-06-19 J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 2 of 30 JC/DOC/325 ISSUE 1.4 REVISIONS PAGE ISSUE RAISED ISSUE DETAILS PAGES AFFECTED 1.0 1.1 1.2 1.3 1.4 1.5 Initial Issue Up issued to address EASA comments EO added Up issued for additional EASA comments Up issued to clarify that the Controlling Pad and associated battery is a carry on item Additional TCDS models added and pod to be installed on either port or starboard strut Typo in description changed All All All All All All J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 3 of 30 JC/DOC/325 ISSUE 1.4 1.0 DESCRIPTION 1.1 This major change is for approval of the aircraft to operate with an XCAM Camera Pod mounted on the port or Starboard Lift Strut. 1.2 The XCAM Camera Pod and attachment interface consist of a Camera Pod and an interface bracket which is attached to either the port or starboard Lift Strut via a clamp arrangement. 1.3 The XCAM is operated by a remote PC/Tablet by the camera operator. The interface is via a single electrical/data harness. The PC/Tablet and associated battery is considered a carry on item. 1.4 This change has been classified Major via the Change / Repair Classification Process Sheet IAW 21.A.91. 2.0 CERTIFICATION BASIS 2.1 The applicable certification basis of the aircraft type is EASA.IM.A.051 (all models), FAA A4EU (all models), A18EU (all models), 3A12(all models), 3A17(all models). However, compliance for this change will be with the latest requirements of CS23 amendment 4 unless stated otherwise. 3.0 COMPLIANCE WITH REQUIREMENTS CS23.25 Weight Limits: (a) Maximum weight. The maximum weight is the highest weight at which compliance with each applicable requirement of CS-23 (other than those complied with at the design landing weight) is shown. The maximum weight must be established The weight of the XCAM and associated equipment is within the provisioned weight limitation of the aircraft. See section 7. CS23.29 Empty Weight and Corresponding Centre of Gravity: (a) The empty weight and corresponding centre of gravity must be determined by weighing the aeroplane with – (1) Fixed ballast; (2) Unusable fuel determined under CS 23.959; and (3) Full operating fluids, including – (i) Oil; (ii) Hydraulic fluid; and (iii) Other fluids required for normal operation of aeroplane systems, except potable water, lavatory precharge water, and water intended for injection in the engines. (b) The condition of the aeroplane at the time of determining empty weight must be one that is well defined and can be easily repeated. The weight of the XCAM is within the provisioned weight limitation of the aircraft. See section 7. CS23.45 Performance General: (a) Unless otherwise prescribed, the performance requirements of this subpart must be met for – (1) Still air and standard atmosphere; (2) Ambient atmospheric conditions, for commuter category aeroplanes, for reciprocating engine- powered aeroplanes of more than 2 722 kg (6 000 lb) maximum weight and for turbine engine- powered aeroplanes. (b) Performance data must be determined over not less than the following ranges of conditions – (1) Aerodrome altitude from sea-level to 3048 m (10 000 ft); J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 4 of 30 JC/DOC/325 ISSUE 1.4 and (2) For reciprocating engine-powered aeroplanes of 2 722 kg (6 000 lb) or less maximum weight, temperatures from standard to 30°C above standard (c) Performance data must be determined with the cowl flaps or other means for controlling the engine cooling air supply in the position used in the cooling tests required by CS 23.1041 to 23.1047. (d) The available propulsive thrust must correspond to engine power, not exceeding the approved power, less – (1) Installation losses; and (2) The power absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and the particular flight condition. (e) The performance as affected by engine power must be based on a relative humidity of – (1) 80% at and below standard temperature; and (2) 34% at and above standard temperature plus 28°C (plus 50°F). Between the two temperatures listed in subparagraphs (e) (1) and (e) (2) the relative humidity must vary linearly. (h) For commuter category aeroplanes, the following also apply: (1) Unless otherwise prescribed, the take-off, en-route, approach and landing configurations for the aeroplane must be selected; (2) The aeroplane configuration may vary with weight, altitude and temperature, to the extent that they are compatible with the operating procedures required by sub-paragraph (h) (3); (3) Unless otherwise prescribed, in determining the critical-engine-inoperative take-off performance, take-off flight path and accelerate-stop distance, changes in the aeroplane’s configuration, speed and power must be made in accordance with procedures established by the applicant for operation in service. (4) Procedures for the execution of discontinued approaches and balked landings associated with the conditions prescribed in CS 23.67 (c) (4) and 23.77 (c) must be established; and (5) The procedures established under sub-paragraphs (h) (3) and (h) (4) must – (i) Be able to be consistently executed by a crew of average skill in atmospheric conditions reasonably expected to be encountered in service; (ii) Use methods or devices that are safe and reliable; and (iii) Include allowances for any reasonably expected time delays in the execution of the procedures. The delta change in performance of the aircraft has been assessed in JC/TR/54. The effect of the pod has been assessed as negligible and has a similar drag to a GOPRO. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.49 Stalling Speed: (a) VSO and VS1 are the stalling speeds or the minimum steady flight speed (CAS) at which the aeroplane is controllable with – (1) For reciprocating engine- powered aeroplanes, engine(s) idling, the throttle(s) closed or at not more than the power necessary for zero thrust at a speed not more than 110% of the stalling speed; and (2) For turbine engine-powered aeroplanes, the propulsive thrust may not be greater than zero at the stalling speed, or, if the resultant thrust has no appreciable effect on the stalling speed, with engine(s) idling and throttle(s) closed; (3) Propeller(s) in the take-off position; (4) The aeroplane in the condition existing in the test in which VSO and VS1 are being used; (5) Centre of gravity in the position which results in the highest value of VSO and VS1; and (6) Weight used when VSO or VS1 are being used as a factor to determine compliance with a required performance standard. (b) VSO and VS1 must be determined by flight tests using the procedure and meeting the flight characteristics specified in CS 23.201. (c) Except as provided in sub-paragraph (d) of this paragraph, VSO at maximum weight must not exceed 113 km/h (61 knots) for – (1) Single-engined aeroplanes; and (2) Twin-engined aeroplanes of 2 722 kg (6 000 lb) or less maximum weight that cannot meet the minimum rate of climb specified in CS 23.67 (a) (1) with the critical engine inoperative. (d) All single-engined aeroplanes, and those twin-engined aeroplanes of 2722 kg (6 000 lb) or less maximum weight, with a VSO of more than 113 km/h (61 knots) at maximum weight that do not meet the requirements of CS 23.67(a)(1), must comply with CS 23.562(d). The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 5 of 30 JC/DOC/325 ISSUE 1.4 CS23.63 Climb: General: (a) Compliance with the requirements of CS 23.65, 23.66, 23.67, 23.69 and 23.77 must be shown – (1) Out of ground effect; and (2) At speeds which are not less than those at which compliance with the powerplant cooling requirements of CS (3) Unless otherwise specified, with one engine inoperative, at a bank angle not exceeding 5 degrees. (b) For normal, utility and aerobatic category reciprocating engine-powered aeroplanes of 2 722 kg (6 000 lb) or less maximum weight, compliance must be shown with CS 23.65 (a), 23.67 (a), where appropriate and CS 23.77 (a) at maximum take-off or landing weight, as appropriate in a standard atmosphere. (c) For normal, utility and aerobatic category reciprocating engined aeroplanes of more than 2 722 kg (6 000 lb) maximum weight and turbine engine-powered aeroplanes in the normal, utility and aerobatic category, compliance must be shown, at weights, as a function of aerodrome altitude and ambient temperature, within the operational limits established for take-off and landing respectively, with – (1) CS 23.65 (b) and 23.67 (b) (1) and (2), where appropriate, for take-off; and (2) CS 23.67 (b) (2), where appropriate, and CS 23.77 (b), for landing. (d) For commuter category aeroplanes, compliance must be shown, at weights as a function of aerodrome altitude and ambient temperature within the operational limits established for take-off and landing respectively, with – (1) CS 23.67 (c) (1), 23.67 (c) (2) and 23.67 (c) (3) for take-off; and (2) CS 23.67 (c) (3), 23.67 (c) (4) and 23.77 (c) for landing. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.65 Climb: All Engines Operating: (a) Each normal, utility and aerobatic category reciprocating engine-powered aeroplane of 2 722 kg (6 000 lb) or less maximum weight must have a steady gradient of climb at sea level of at least 8·3% for landplanes or 6·7% for seaplanes and amphibians with – (1) Not more than maximum continuous power on each engine; (2) The landing gear retracted; (3) The wing flaps in the take-off position(s); and (4) A climb speed not less than the greater of 1·1 VMC and 1·2 VS1 for twin engined aeroplanes and not less than 1·2 VS1 for single-engined aeroplanes. (b) Each normal, utility and aerobatic category reciprocating engine-powered aeroplanes of more than 2 722 kg (6 000 lb) maximum weight and turbine engine-powered aeroplanes in the normal, utility and aerobatic category must have a steady gradient of climb after take-off of at least 4% with – (1) Take-off power on each engine; (2) The landing gear extended except that, if the landing gear can be retracted in not more than 7 seconds, it may be assumed to be retracted; (3) The wing flaps in the take-off position(s); and (4) A climb speed as specified in CS 23.65 (a) (4). The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.69 En-Route Climb/Descent (a) All engines operating The steady gradient and rate of climb must be determined at each weight, altitude and ambient temperature within the operational limits established by the applicant with – (1) Not more than maximum continuous power on each engine; (2) The landing gear retracted; (3) The wing flaps retracted; and (4) A climb speed not less than 1·3 VS1. (b) One-engine-inoperative The steady gradient and rate of climb/descent must be determined at each weight, altitude and ambient temperature within the operational limits established by the applicant with – (1) The critical engine inoperative and its propeller in the minimum drag position; (2) The remaining engine at not more than maximum continuous power; (3) The landing gear retracted; (4) The wing flaps retracted; and (5) A climb speed not less than 1·2 VS1. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 6 of 30 JC/DOC/325 ISSUE 1.4 The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS 23.71 Glide (Single-Engined Aeroplanes): The maximum horizontal distance travelled in still air, in km per 1000 m (nautical miles per 1 000 ft) of altitude lost in a glide, and the speed necessary to achieve this, must be determined with the engine inoperative and its propeller in the minimum drag position, landing gear and wing flaps in the most favourable available position. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.73 Reference Landing Approach Speed: (a) For normal, utility and aerobatic category reciprocating engine-powered aeroplanes of 2 722 kg (6 000 lb) or less maximum weight, the reference landing approach speed, VREF, must not be less than the greater of VMC, determined under CS 23.149 (b) with the wing flaps in the most extended take-off setting, and 1·3 VSO. (b) For normal, utility and aerobatic category reciprocating engine-powered aeroplanes of more than 2 722 kg (6 000 lb) maximum weight and turbine engine-powered aeroplanes in the normal, utility and aerobatic category, the reference landing approach speed, VREF, must not be less than the greater of VMC, determined under CS 23.149 (c), and 1·3 VS0. (c) For commuter category aeroplanes, the reference landing approach speed, VREF, must not be less than the greater of 1·05 VMC, determined under CS 23.149 (c), and 1·3 VSO. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS 23.75 Landing Distance: The horizontal distance necessary to land and come to a complete stop from a point 15 m (50 ft) above the landing surface must be determined, for standard temperatures at each weight and altitude within the operational limits established for landing, as follows: (a) A steady approach at not less than VREF, determined in accordance with CS 23.73 (a), (b) or (c) as appropriate, must be maintained down to 15 m (50 ft) height. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.77 Balked Landing: (a) Each normal, utility and aerobatic category reciprocating engine- powered aeroplane of 2 722 kg (6 000 lb) or less maximum weight must be able to maintain a steady gradient of climb at sea-level of at least 3·3% with – (1) Take-off power on each engine; (2) The landing gear extended; (3) The wing flaps in the landing position, except that if the flaps may safely be retracted in two seconds or less without loss of altitude and without sudden changes of angle of attack, they may be retracted; and (4) A climb speed equal to VREF, as defined in CS 23.73 (a). (b) For normal, utility and aerobatic category each reciprocating engine- powered aeroplane of more than 2 722 kg (6 000 lb) maximum weight and turbine engine- powered aeroplanes in the normal, utility and aerobatic category, the steady gradient of climb must not be less than 2·5% with – (1) Not more than the power or thrust that is available 8 seconds after initiation of movement of the power controls from the minimum flight-idle position; (2) The landing gear extended; (3) The wing flaps in the landing position; and (4) A climb speed equal to VREF, as defined in CS 23.73 (b). (c) For each commuter category aeroplane, the steady gradient of climb must not be less than 3·2% with – (1) Not more than the power that is J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 7 of 30 JC/DOC/325 ISSUE 1.4 available 8 seconds after initiation of movement of the power controls from the minimum flight idle position; (2) Landing gear extended; (3) Wing flaps in the landing position; and (4) A climb speed equal to VREF, as defined in CS 23.73 (c). The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.141 Flight Characteristics General: The aeroplane must meet the requirements of CS 23.143 to 23.253 at all practical loading conditions and all operating altitudes, not exceeding the maximum operating altitude established under CS 23.1527, for which certification has been requested, without requiring exceptional piloting skill, alertness or strength. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on the flight characteristics. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.143 Control and Manoeuvrability General: The aeroplane must be safely controllable and manoeuvrable during all flight phases including (1) Take-off; (2) Climb; (3) Level flight; (4) Descent; (5) Go-around; and (6) Landing (power on and power off) with the wing flaps extended and retracted. (b) It must be possible to make a smooth transition from one flight condition to another (including turns and slips) without danger of exceeding the limit load factor, under any probable operating condition, (including, for multi-engined aeroplanes, those conditions normally encountered in the sudden failure of any engine). (c) If marginal conditions exist with regard to required pilot strength, the control forces required must be determined by quantitative tests. In no case may the control forces under the conditions specified in sub-paragraphs (a) and (b), exceed those prescribed in the following table: The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on control and manoeuvrability. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.145 Longitudinal Control (a) With the aeroplane as nearly as possible in trim at 1·3 VS1, it must be possible, at speeds below the trim speed, to pitch the nose downward so that the rate of increase in airspeed allows prompt acceleration to the trim speed with – (1) Maximum continuous power on each engine; (2) Power off; and (3) Wing flaps and landing gear – (i) J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 8 of 30 JC/DOC/325 ISSUE 1.4 Retracted; and (ii) Extended. (b) It must be possible to carry out the following manoeuvres without requiring the application of single handed control forces exceeding those specified in CS 23.143 (c), unless otherwise stated. The trimming controls must not be adjusted during the manoeuvres: (1) With landing gear extended and flaps retracted and the aeroplane as nearly as possible in trim at 1·4 VS1, extend the flaps as rapidly as possible and allow the airspeed to transition from 1·4 VS1 to 1·4 VS0, with – (i) Power off; and (ii) Power necessary to maintain level flight in the initial condition. (2) With landing gear and flaps extended, power off and the aeroplane as nearly as possible in trim at 1·3 VSO, quickly apply take-off power and retract the flaps as rapidly as possible to the recommended go-around setting and allow the airspeed to transition from 1·3 VSO to 1·3 VS1. Retract the gear when a positive rate of climb is established. (3) With landing gear and flaps extended, power for and in level flight at 1·1 VSO and the aeroplane as nearly as possible in trim, it must be possible to maintain approximately level flight while retracting the flaps as rapidly as possible with simultaneous application of not more than maximum continuous power. If gated flap positions are provided, the flap retraction may be demonstrated in stages with power and trim reset for level flight at 1·1 VS1 in the initial configuration for each stage – (i) From the fully extended position to the most extended gated position; (ii) Between intermediate gated positions, if applicable; and (iii) From the least extended gated position to the fully retracted position. (4) With power off, flaps and landing gear retracted and the aeroplane as nearly as possible in trim at 1·4 VS1, apply take-off power rapidly while maintaining the same airspeed. (5) With power off, landing gear and flaps extended and the aeroplane as nearly as possible in trim at VREF, obtain and maintain airspeeds between 1·1 VS0 and either 1·7 VS0 or VFE, whichever is lower, without requiring the application of two-handed control forces exceeding those specified in CS 23.143 (c). (6) With maximum take-off power, landing gear retracted, flaps in the take-off position and the aeroplane as nearly as possible in trim at VFE appropriate to the take-off flap position, retract the flaps as rapidly as possible while maintaining speed constant. (c) At speeds above VMO/MMO and up to the maximum speed shown under CS 23.251, a manoeuvring capability of 1·5g must be demonstrated to provide a margin to recover from upset or inadvertent speed increase. (d) It must be possible, with a pilot control force of not more than 44·5 N (10 lbf), to maintain a speed of not more than VREF during a power-off glide with landing gear and wing flaps extended. (e) By using normal flight and power controls, except as otherwise noted in subparagraphs (e) (1) and (e) (2) , it must be possible to establish a zero rate of descent at an attitude suitable for a controlled landing without exceeding the operational and structural limitations of the aeroplane, as follows: (1) For single-engined and twin-engined aeroplanes, without the use of the primary longitudinal control system; (2) For twin- engined aeroplanes; (i) Without the use of the primary directional control; and (ii) If a single failure of any one connecting or transmitting link would affect both the longitudinal and directional primary control system, without the primary longitudinal and directional control system. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on longitudinal control. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.147 Directional and Lateral Control: (a) For each twin-engined aeroplane, it must be possible, while holding the wings level within 5°, to make sudden changes in heading safely in both directions. This must be shown at 1·4 VS1 with heading changes up to 15° (except that the heading change at which the rudder force corresponds to the limits specified in CS 23.143 need not be exceeded), with the – (1) Critical engine inoperative and its propeller in the minimum drag position; (2) Remaining engine at maximum continuous power; (3) Landing gear – (i) Retracted; and (ii) Extended; and (4) Flaps retracted. (b) For each twin-engined aeroplane, it must be possible to regain full control of the aeroplane without exceeding a bank angle of 45°, reaching a dangerous attitude or encountering dangerous characteristics, in the J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 9 of 30 JC/DOC/325 ISSUE 1.4 event of a sudden and complete failure of the critical engine, making allowance for a delay of 2 seconds in the initiation of recovery action appropriate to the situation, with the aeroplane initially in trim, in the following conditions – (1) Maximum continuous power on each engine; (2) Wing flaps retracted; (3) Landing gear retracted; (4) Speed equal to that at which compliance with CS 23.69 (a) has been shown; (5) All propeller controls in the position in which compliance with CS 23.69 (a) has been shown. (c) For all aeroplanes, it must be shown that the aeroplane is safely controllable without the use of the primary lateral control system in any all engine configuration(s) and at any speed or altitude within the approved operating envelope. It must also be shown that the aeroplane’s flight characteristics are not impaired below a level needed to permit continued safe flight and the ability to maintain attitudes suitable for a controlled landing without exceeding the operational and structural limitations of the aeroplane. If a single failure of any one connecting or transmitting link in the lateral control system would also cause the loss of additional control system(s), the above requirement is equally applicable with those additional systems also assumed to be inoperative. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on directional and lateral control. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS 23.151 Aerobatic Manoeuvres: Each aerobatic and utility category aeroplane must be able to perform safely the aerobatic manoeuvres for which certification is requested. Safe entry speeds for these manoeuvres must be determined. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on aerobatic manoeuvres. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.153 Control During Landings: It must be possible, while in the landing configuration, to safely complete a landing without exceeding the one-hand control force limits specified in CS 23.143 (c) following an approach to land – (a) At a speed of VREF –9.3 km/h (5 knots); (b) With the aeroplane in trim, or as nearly as possible in trim and without the trimming control being moved throughout the manoeuvre; (c) At an approach gradient equal to the steepest used in the landing distance demonstration of CS 23.75; (d) With only those power changes, if any, which would be made when landing normally from an approach at VREF. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on control during landings. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.157 Rate of Roll: a) Take–off. It must be possible, using a favourable combination of controls, to roll the aeroplane from a steady 30° banked turn through an angle of 60°, so as to reverse the direction of the turn within – (1) For an aeroplane of 2 722 kg (6 000 lb) or less maximum weight, 5 seconds from initiation of roll; and (2) For aeroplanes of over 2 722 kg (6 000 lb) maximum weight, but not more than 10 seconds, where W is the weight in kg, but not more than 10 seconds, where W is the weight in lb.) (b) The requirement of sub-paragraph (a) must be met when rolling the aeroplane in each direction in the following conditions – (1) Flaps in the take-off position; (2) Landing gear retracted; (3) For a single-engined aeroplane, at maximum take-off power and for a twin engined aeroplane, with the critical engine inoperative, the propeller in the minimum drag position and the remaining engine at maximum take-off power; and (4) The aeroplane trimmed at a speed equal to the greater of 1·2 VS1 or 1·1 VMC or as nearly as possible in trim for straight flight. (c) Approach. It must be possible using a J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 10 of 30 JC/DOC/325 ISSUE 1.4 favourable combination of controls, to roll the aeroplane from a steady 30° banked turn through an angle of 60°, so as to reverse the direction of the turn within – (1) For an aeroplane of 2 722 kg (6 000 lb) or less maximum weight, 4 seconds from initiation of roll; and (2) For and aeroplane of over 2 722 kg (6 000 lb) maximum weight, but not more than 7 seconds where W is weight in kg. 2 200 W 2 800 but not more than 7 seconds where W is weight in lb.) (d) The requirement of sub-paragraph (c) must be met when rolling the aeroplane in each direction in the following conditions – (1) Flaps in the landing position(s); (2) Landing gear extended; (3) All engines operating at the power for a 3° approach; and (4) The aeroplane trimmed at VREF. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on rate of roll. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.161 Trim: (a) General. Each aeroplane must meet the trim requirements after being trimmed and without further pressure upon, or movement of, the primary controls or their corresponding trim controls by the pilot or the automatic pilot. In addition, it must be possible, in other conditions of loading, configuration, speed and power to ensure that the pilot will not be unduly fatigued or distracted by the need to apply residual control forces exceeding those for prolonged application of CS 23.143 (c). This applies in normal operation of the aeroplane and, if applicable, to those conditions associated with the failure of one engine for which performance characteristics are established. (b) Lateral and directional trim. The aeroplane must maintain lateral and directional trim in level flight with the landing gear and wing flaps retracted as follows: (1) For normal, utility and aerobatic category aeroplanes, at a speed of 0·9 VH, VC or VMO/MMO, whichever is lowest; and(2) For commuter category aeroplanes, at all speeds from 1·4 VSI to the lesser of VH or VMO/MMO. (c) Longitudinal trim. The aeroplane must maintain longitudinal trim under each of the following conditions: (1) A climb with; (i) Take-off power, landing gear retracted, wing flaps in the take-off position(s), at the speeds used in determining the climb performance required by CS 23.65; and (ii) Maximum continuous power at the speeds and in the configuration used in determining the climb performance required by CS 23.69 (a). (2) Level flight at all speeds from the lesser of VH and either VNO or VMO/MMO (as appropriate), to 1·4 VS1, with the landing gear and flaps retracted. (3) A descent at VNO or VMO/MMO, whichever is applicable, with power off and with the landing gear and flaps retracted. (4) Approach with landing gear extended and with – (i) A 3° angle of descent, with flaps retracted and at a speed of 1·4 VS1; (ii) A 3° angle of descent, flaps in the landing position(s) at VREF; and (iii) An approach gradient equal to the steepest used in the landing distance demonstrations of CS 23.75, flaps in the landing position(s) at VREF. (d) In addition, each twin-engined aeroplane must maintain longitudinal and directional trim and the lateral control force must not exceed 22 N (5 lbf), at the speed used in complying with CS 23.67 (a) or (b) (2) or (c) (3) as appropriate, with – (1) The critical engine in-operative and its propeller in the minimum drag position; (2) The remaining engine at maximum continuous power; (3) The landing gear retracted; (4) The wing flaps retracted; and (5) An angle of bank of not more than 5°. (e) In addition, each commuter category aeroplane for which, in the determination of the take-off path in accordance with CS 23.57, the climb in the take-off configuration at V2 extends beyond 122 m (400 ft) above the take-off surface, it must be possible to reduce the longitudinal and lateral control forces to 44·5 N (10 lbf) and 22 N (5 lbf) respectively and the directional control force must not exceed 222 N (50 lbf) at V2 with – (1) The critical engine inoperative and its propeller in the minimum drag position; (2) The remaining engine at take-off power; (3) Landing gear retracted; (4) Wing flaps in the take- off position(s); and (5) An angle of bank not exceeding 5°. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 11 of 30 JC/DOC/325 ISSUE 1.4 installation of the XCAM has a negligible impact on trim. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.177 Static Directional and Lateral Stability: (a) The static directional stability, as shown by the tendency to recover from a wings level sideslip with the rudder free, must be positive for any landing gear and flap position appropriate to the take-off, climb, cruise, approach and landing configurations. This must be shown with symmetrical power up to maximum continuous power and at speeds from 1·2 VS1 up to maximum allowable speed for the condition being investigated. The angle of sideslip for these tests must be appropriate to the type of aeroplane. At larger angles of sideslip up to that at which full rudder is used or a control force limit in CS 23.143 is reached, whichever occurs first, and at speeds from 1·2 VS1 to Vo the rudder pedal force must not reverse. (b) The static lateral stability, as shown by the tendency to raise the low wing in a sideslip, must be positive for all landing gear and flap positions. This must be shown with symmetrical power up to 75% of maximum continuous power at speeds above 1·2 VS1 in the take-off configuration(s) and at speeds above 1·3 VS1 in other configurations, up to the maximum allowable speed for the configuration being investigated, in the take-off, climb, cruise and approach configurations. For the landing configuration, the power must be up to that necessary to maintain a 3° angle of descent in coordinated flight. The static lateral stability must not be negative at 1·2 VS1 in the take-off configuration, or at 1·3 VS1 in other configurations. The angle of sideslip for these tests must be appropriate to the type of aeroplane but in no case may the constant heading sideslip angle be less than that obtainable with 10° bank, or if less, the maximum bank angle obtainable with full rudder deflection or 667 N (150 lbf) rudder force.(c) Sub-paragraph (b) does not apply to aerobatic category aeroplanes certificated for inverted flight. (d) In straight, steady sideslips at 1·2 VS1 for any landing gear and flap positions and for any symmetrical power conditions up to 50% of maximum continuous power, the aileron and rudder control movements and forces must increase steadily (but not necessarily in constant proportion) as the angle of sideslip is increased up to the maximum appropriate to the type of aeroplane. At larger sideslip angles up to the angle at which full rudder or aileron control is used or a control force limit contained in CS 23.143 is reached, the aileron and rudder control movements and forces must not reverse as the angle of sideslip is increased. Rapid entry into, or recovery from, a maximum sideslip considered appropriate for the aeroplane must not result in uncontrollable flight characteristics. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on directional and lateral stability. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.181 Dynamic Stability: (a) Any short period oscillation not including combined lateral- directional oscillations occurring between the stalling speed and the maximum allowable speed appropriate to the configuration of the aeroplane must be heavily damped with the primary controls – (1) Free; and (2) In a fixed position, except when compliance with CS 23.672 is shown. (b) Any combined lateral–directional oscillations (“Dutch roll”) occurring between the stalling speed and the maximum allowable speed appropriate to the configuration of the aeroplane must be damped to XXX amplitude in 7 cycles with the primary controls – (1) Free; and (2) In a fixed position, except when compliance with CS 23.672 is shown. (c) Any long-period oscillation of the flight path (phugoid) must not be so unstable as to cause an unacceptable increase in pilot workload or otherwise endanger the aeroplane. When, in the conditions of CS 23.175, the longitudinal control force required to maintain speeds differing from the trimmed speed by at least plus or minus 15% is suddenly released, the response of the aeroplane must not exhibit any dangerous characteristics nor be excessive in relation to the magnitude of the control force released. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 12 of 30 JC/DOC/325 ISSUE 1.4 The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on the dynamic stability. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.201 Wing Level Stall: (a) It must be possible to produce and to correct roll by unreversed use of the rolling control and to produce and to correct yaw by unreversed use of the directional control, up to the time the aeroplane stalls. (b) The wings level stall characteristics must be demonstrated in flight as follows. Starting from a speed at least 18.5 km/h (10 knots) above the stall speed, the elevator control must be pulled back so that the rate of speed reduction will not exceed 1.9 km/h (one knot) per second until a stall is produced, as shown by either – (1) An uncontrollable downward pitching motion of the aeroplane; or (2) A downward pitching motion of the aeroplane which results from the activation of a device (e.g. stick pusher); or (3) The control reaching the stop. (c) Normal use of elevator control for recovery is allowed after the downward pitching motion of (b) (1) or (b) (2) has unmistakably been produced, or after the control has been held against the stop for not less than the longer of 2 seconds or the time employed in the minimum steady flight speed determination of CS 23.49. (d) During the entry into and the recovery from the manoeuvre, it must be possible to prevent more than 15° of roll or yaw by the normal use of controls. (e) Compliance with the requirements must be shown under the following conditions: (1) Wing flaps. Retracted, fully extended and each intermediate normal operating position; (2) Landing gear. Retracted and extended; (3) Cowl flaps. Appropriate to configuration; (4) Power (i) Power off; and (ii) 75% maximum continuous power. If the power-to-weight ratio at 75% of maximum continuous power results in extreme nose-up attitudes, the test may be carried out with the power required for level flight in the landing configuration at maximum landing weight and a speed of 1·4 VS0, but the power may not be less than 50% maximum continuous power. (5) Trim. The aeroplane trimmed at a speed as near 1·5 VS1 as practicable. (6) Propeller. Full increase rpm position for the power off condition. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on wing level stall. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.203 Turning Flight and Accelerated Turning Stalls: Turning flight and accelerated turning stalls must be demonstrated in tests as follows: (a) Establish and maintain a co- ordinated turn in a 30° bank. Reduce speed by steadily and progressively tightening the turn with the elevator until the aeroplane is stalled, as defined in CS 23.201 (b). The rate of speed reduction must be constant, and – (1) For a turning flight stall, may not exceed 1.9 km/h (one knot) per second; and (2) For an accelerated turning stall, be 5.6 to 9.3 km/h (3 to 5 knots) per second with steadily increasing normal acceleration. (b) After the aeroplane has stalled, as defined in CS 23.201 (b) it must be possible to regain level flight by normal use of the flight controls but without increasing power and without – (1) Excessive loss of altitude; (2) Undue pitch-up; (3) Uncontrollable tendency to spin; (4) Exceeding a bank angle of 60° in the original direction of the turn or 30° in the opposite direction, in the case of turning flight stalls; (5) Exceeding a bank angle of 90° in the original direction of the turn or 60° in the opposite direction, in the case of accelerated turning stalls; and (6) Exceeding the maximum permissible speed or allowable limit load factor. (c) Compliance with the requirements must be shown under the following conditions: (1) Wing flaps. Retracted, fully extended and each intermediate normal operating position; (2) Landing gear. Retracted and extended; (3) Cowl flaps. Appropriate to configuration; (4) Power (i) Power off; and (ii) 75% maximum continuous power. If the power-to-weight ratio at 75% of maximum continuous power results in extreme nose-up attitudes, the test may be carried out with the power required for level flight in the landing configuration at maximum landing weight and a speed of 1·4 VS0, but the power may J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 13 of 30 JC/DOC/325 ISSUE 1.4 not be less than 50% maximum continuous power. (5) Trim. The aeroplane trimmed at a speed as near 1·5 VS1 as practicable. (6) Propeller. Full increase rpm position for the power off condition. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on the turning flight and accelerated turning stalls. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.207 Stall Warning: (a) There must be a clear and distinctive stall warning, with the flaps and landing gear in any normal position, in straight and turning flight. (b) The stall warning may be furnished either through the inherent aerodynamic qualities of the aeroplane or by a device that will give clearly distinguishable indications under expected conditions of flight. However, a visual stall warning device that requires the attention of the crew within the cockpit is not acceptable by itself. (c) During the stall tests required by CS 23.201 (b) and CS 23.203 (a) (1), the stall warning must begin at a speed exceeding the stalling speed by a margin of not less than 9.3 km/h (5 knots) and must continue until the stall occurs. (d) When following the procedures of CS 23.1585, the stall warning must not occur during a take-off with all engines operating, a take-off continued with one engine inoperative or during an approach to landing. (e) During the stall tests required by CS 23.203 (a) (2), the stall warning must begin sufficiently in advance of the stall for the stall to be averted by pilot action taken after the stall warning first occurs. (f) For aerobatic category aeroplanes, an artificial stall warning may be mutable, provided that it is armed automatically during take-off and re-armed automatically in the approach configuration. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. The installation of the XCAM has a negligible impact on the stall warning. See JC/TR/54 and JC/FMS/4584 as defined in section 11.See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.251 Vibration and Buffeting: There must be no vibration or buffeting severe enough to result in structural damage and each part of the aeroplane must be free from excessive vibration, under any appropriate speed and power conditions up to at least the minimum value of VD allowed in CS 23.335. In addition there must be no buffeting in any normal flight condition severe enough to interfere with the satisfactory control of the aeroplane or cause excessive fatigue to the flight crew. Stall warning buffeting within these limits is allowable. The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. An advisory post installation confirmatory flight has been defined in the EO to ensure no vibration or buffeting is present prior to release to the operator. See JC/TR/54, JC/EO/4584-1 and JC/FMS/4584 as defined in section 11. CS23.301 Loads: (a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced aerodynamic and inertia loads are considered small and have been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 14 of 30 JC/DOC/325 ISSUE 1.4 CS23.303 Factor of Safety: Unless otherwise specified, a factor of safety of 1 ∙5 must be used. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced aerodynamic and inertia loads are considered small and have been assessed in JC/TR/54. The factor of safety of 1.5 is not affected. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS25.305 Strength and Deformation: (a) The structure must be able to support limit loads without detrimental, permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation. (b) The structure must be able to support ultimate loads without failure for at least three seconds, except local failures or structural instabilities between limit and ultimate load are acceptable only if the structure can sustain the required ultimate load for at least three seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the three second limit does not apply. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced aerodynamic and inertia loads are considered small and have been assessed in JC/TR/54. The strength and deformation is not affected. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.307 Proof of Structures: (a) Compliance with the strength and deformation requirements of CS 23.305 must be shown for each critical load condition. Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable. In other cases, substantiating load tests must be made. Dynamic tests, including structural flight tests, are acceptable if the design load conditions have been simulated.(b) Certain parts of the structure must be tested as specified in Subpart D of CS-23. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced aerodynamic and inertia loads are considered small and have been assessed in JC/TR/54. The proof of structure is not affected. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS 23.321 Flight Loads General: (a) Flight load factors represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the aeroplane) to the weight of the aeroplane. A positive flight load factor is one in which the aerodynamic force acts upward, with respect to the aeroplane. (b) Compliance with the flight load requirements of this subpart must be shown – (1) At each critical altitude within the range in which the aeroplane may be expected to operate; (2) At each weight from the design minimum weight to the design maximum weight; and (3) For each required altitude and weight, for any practicable distribution of disposable load within the operating (c) When significant the effects of compressibility must be taken into account. The affects of flight loads on the XCAM installation have been assessed in JC/TR/54. See JC/TR/54 as defined in section 11. CS23.333 Flight Envelope: (a) General. Compliance with the strength requirements of this subpart must be shown at any combination of airspeed and load factor on and within the boundaries of a flight envelope (similar to the one in sub-paragraph (d) ) that represents the envelope of the flight loading conditions specified by the manoeuvring and gust criteria of sub- paragraphs (b) and (c) respectively. (b) Manoeuvring envelope. Except where limited by maximum (static) lift coefficients, the aeroplane is assumed to be subjected to symmetrical manoeuvres resulting in the following limit load factors: (1) The positive manoeuvring load factor specified in CS 23.337 at speeds up to VD; (2) The negative manoeuvring load factor specified in CS 23.337 at VC; and (3) Factors varying linearly with speed from the specified J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 15 of 30 JC/DOC/325 ISSUE 1.4 value at VC to 0·0 at VD for the normal and commuter category, and -1·0 at VD for the aerobatic and utility categories. (c) Gust envelope (1) The aeroplane is assumed to be subjected to symmetrical vertical gusts in level flight. The resulting limit load factors must correspond to the conditions determined as follows: (i) Positive (up) and negative (down) gusts of 50 fps at VC must be considered at altitudes between sea level and 6096 m (20 000 ft). The gust velocity may be reduced linearly from 50 fps at 6096 m (20 000 ft) to 25 fps at 15240 m (50 000 ft); and (ii) Positive and negative gusts of 25 fps at VD must be considered at altitudes between sea level and 6096 m (20 000 ft). The gust velocity may be reduced linearly from 25 fps at 6096 m (20 000 ft) to 12·5 fps at 15240 m (50 000 ft). (iii) In addition, for commuter category aeroplanes, positive (up) and negative (down) rough air gusts of 66 fps at VB must be considered at altitudes between sea level and 6096 m (20 000 ft). The gust velocity may be reduced linearly from 66 fps at 6096 m (20 000 ft) to 38 fps at 15240 m (50 000 ft). (2) The following assumptions must be made: (i) The shape of the gust is – s = Distance penetrated into gust (ft.); C= Mean geometric chord of wing (ft.); and Ude = Derived gust velocity referred to in sub-paragraph (1) linearly with speed between VC and VD. (ii) Gust load factors vary linearly with speed between VC and VD. The VNE with the XCAM installed has not been restricted. The affects of flight loads on the XCAM have been assessed under JC/TR/54. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.335 Design Airspeed: Except as provided in sub-paragraph (a) (4), the selected design airspeeds are equivalent airspeeds (EAS). The VNE with the XCAM installed has not been restricted see JC/FMS/4584. The affects of flight loads on the XCAM have been assessed under JC/TR/54. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS23.471 Ground Loads General: The limit ground loads specified in this subpart are considered to be external loads and inertia forces that act upon an aeroplane structure. In each specified ground load condition, the external reactions must be placed in equilibrium with the linear and angular inertia forces in a rational or conservative manner. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced grounds loads are considered small and have been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.561 Emergency Landing General: a) The aeroplane, although it may be damaged in emergency landing conditions, must protect each occupant under those conditions. (b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury when – (1) Proper use is made of seats, safety belts and shoulder harnesses provided for in the design; (2) The occupant experiences the static inertia loads corresponding to the following ultimate load factors: (i) Upward, 3·0g for normal, utility, and commuter category aeroplanes, or 4·5g for aerobatic category aeroplanes; (ii) Forward, 9·0g; (iii) Sideward, 1·5g; and (iv) Downward, 6·0g when certification to the emergency exit provisions of sub-paragraph 23.807(d)(4) is requested; and (3) The items of mass within the cabin, that could injure an occupant, experience the static inertia loads corresponding to the following ultimate load factors: (i) Upward, 3·0g; (ii) Forward, 18·0g; and (iii) Sideward, 4·5g. (c) Each aeroplane with retractable landing gear must be designed to protect each occupant in a landing – (1) With the wheels retracted; (2) With moderate descent velocity; And (3) Assuming, in the absence of a more rational analysis – (i) A downward ultimate inertia force of 3g; and (ii) A coefficient of friction of 0·5 at the ground. (d) If it is not established that a turnover is unlikely during an emergency landing, the structure must be designed to protect the occupants in a complete turnover as follows: (1) The likelihood of a turnover may be J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 16 of 30 JC/DOC/325 ISSUE 1.4 shown by an analysis assuming the following conditions: (i) The most adverse combination of weight and centre of gravity position; (ii) Longitudinal load factor of 9·0g; (iii) Vertical load factor of 1·0g; And (iv) For aeroplanes with tricycle landing gear, the nose wheel strut failed with the nose contacting the ground. (2) For determining the loads to be applied to the inverted aeroplane after a turnover, an upward ultimate inertia load factor of 3·0g and a coefficient of friction with the ground of 0·5 must be used. (e) Except as provided in CS 23.787 (c) the supporting structure must be designed to restrain, under loads up to those specified in subparagraph (b) (3) , each item of mass that could injure an occupant if it came loose in a minor crash landing. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The PC/Tablet (which controls the XCAM) and Power Unit are stowed for Taxi, Take Off and Landing. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.571 Fatigue Evaluation Metallic Pressurised Structures: For normal, utility, and aerobatic category aeroplanes, the strength, detail design, and fabrication of the metallic structure of the pressure cabin must be evaluated under one of the following:- (a) A fatigue strength investigation in which the structure is shown by tests, or by analysis supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected in service; or (b) A fail safe strength investigation, in which it is shown by analysis, tests, or both that catastrophic failure of the structure is not probable after fatigue failure, or obvious partial failure, of a principal structural element, and that the remaining structures are able to withstand a static ultimate load factor of 75 percent of the limit load factor at Vc, considering the combined effects of normal operating pressures, expected external aerodynamic pressures, and flight loads. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise considered (c) The damage tolerance evaluation of CS 23.573(b). Not applicable as the aircraft is not pressurised. See TCDS as defined in section 2. CS23.572 Metallic Wing, Empennage and Associated Structures (See AMC to 23.571 and 23.572) (a) For normal, utility, and aerobatic category aeroplanes, the strength, detail design, and fabrication of those parts of the airframe structure whose failure would be catastrophic must be evaluated under one of the following unless it is shown that the structure, operating stress level, materials and expected uses are comparable, from a fatigue standpoint, to a similar design that has had extensive satisfactory service experience: (1) A fatigue strength investigation in which the structure is shown by tests, or by analysis supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected in service; or (2) A fail-safe strength investigation in which it is shown by analysis, tests, or both, that catastrophic failure of the structure is not probable after fatigue failure, or obvious partial failure, of a principal structural element, and that the remaining structure is able to withstand a static ultimate load factor of 75 percent of the critical limit load factor at Vc. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise considered. (3) The damage tolerance evaluation of CS23.573(b) b) Each evaluation required by this paragraph must:- (1) Include typical loading spectra (e.g. taxi, ground-air-ground cycles manoeuvre, gust); (2) Account for any significant effects due to the mutual influence of aerodynamic surfaces; and (3) Consider any significant effects from propeller slipstream loading, and buffet from vortex impingements. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The additional induced aerodynamic and inertia loads are considered small and have been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 17 of 30 JC/DOC/325 ISSUE 1.4 CS23.573 Damage Tolerance and Fatigue Evaluation: (b) Metallic airframe structure. If the applicant elects to use CS 23.571(c) or CS 23.572(a)(3), then the damage tolerance evaluation must include a determination of the probable locations and modes of damage due to fatigue, corrosion, or accidental damage. The determination must be by analysis supported by test evidence and, if available, service experience. Damage at multiple sites due to fatigue must be included where the design is such that this type of damage can be expected to occur. The evaluation must incorporate repeated load and static analyses supported by test evidence. The extent of damage for residual strength evaluation at any time within the operational life of the aeroplane must be consistent with the initial detectability and subsequent growth under repeated loads. The residual strength evaluation must show that the remaining structure is able to withstand critical limit flight loads, considered as ultimate, with the extent of detectable damage consistent with the results of the damage tolerance evaluations. For pressurised cabins, the following load must be withstood: (1) The normal operating differential pressure combined with the expected external aerodynamic pressures applied simultaneously with the flight loading conditions specified in this subpart, an (2) The expected external aerodynamic pressures in 1g flight combine with a cabin differential pressure equal to 1. times the normal operating differential pressure without any other load. Not affected by this change. The XCAM is attached to the port or starboard Lift Strut via a clamping arrangement. The XCAM and clamping arrangement are installed within the weight limitations of fully provisioned structure. The XCAM and clamping arrangement have been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS 23.575 Inspections and other Procedures Each inspection or other procedure, based on an evaluation required by CS paragraphs 23.571, 23.572, 23.573 or 23.574, must be established to prevent catastrophic failure and must be included in the limitations section of the instructions for continued airworthiness required by CS 23.1529. Not affected by this change. The XCAM and clamping arrangement are installed within the weight limitations of fully provisioned structure. No new inspections of the aircraft structure are required. See JC/FMS/4584 as defined in section 11. CS23.601 Design and Construction General: The suitability of each questionable design detail and part having an important bearing on safety in operations must be established by tests. Not affected by this change. No new methods of fabrication are introduced and have no details that are hazardous or unreliable. The installation has been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.603 Materials and Workmanship: (a) The suitability and durability of materials used for parts, the failure of which could adversely affect safety, must – (1) Be established by experience or tests; (2) Meet approved specifications that ensure their having the strength and other properties assumed in the design data; and (3) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service. (b) Workmanship must be of a high standard. Not affected by this change. The XCAM installation has been assessed in JC/TR/54 and no new materials, materials that are not approved or methods of fabrication are introduced. All work and protection of structure is carried out IAW standard aircraft practices. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 18 of 30 JC/DOC/325 ISSUE 1.4 CS23.605 Fabrication Methods: (a) The methods of fabrication used must produce consistently sound structures. If a fabrication process (such as gluing, spot welding, or heat treating) requires close control to reach this objective, the process must be performed under an approved process specification. (b) Each new aircraft fabrication method must be substantiated by a test programme. No new methods of fabrication are introduced and have no details that are hazardous or unreliable. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.607 Fasteners: a) Each removable fastener must incorporate two retaining devices if the loss of such fastener would preclude continued safe flight and landing. (b) Fasteners and their locking devices must not be adversely affected by the environmental conditions associated with the particular installation. (c) 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 locking device. Not affected by this change. The existing structure and fasteners are defined on the installation drawing and assessed in JC/TR/54. See ECN(s)/Drawing(s) as defined in section 11. CS23.609 Protection of Structure: (a) Be suitably protected against deterioration or loss of strength in service due to any cause, including – (1) Weathering; (2) Corrosion; and (3) Abrasion; and (b) Have adequate provisions for ventilation and drainage. Not affected by this change. The installation of the XCAM installation has been protected IAW standard aircraft practices against deterioration or loss of strength by weathering, corrosion, and abrasion. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.611 Accessibility Provisions: For each part that requires maintenance, inspection, or other servicing, appropriate means must be incorporated into the aircraft design to allow such servicing to be accomplished. The removal and installation of the XCAM does not affect access, inspection, replacement, adjustment or lubrication of parts. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.613 Material Strength Properties and Design Values: (a) Material strength properties must be based on enough tests of material meeting specifications to establish design values on a statistical basis. (b) The design values must be chosen to minimise the probability of structural failure due to material variability. Except as provided in subparagraph (e) , compliance with this paragraph must be shown by selecting design values that assure material strength with the following probability: (1) Where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component ; 99% probability with 95% confidence. (2) For redundant structure, in which the failure of individual elements would result in applied loads being safely distributed to other load carrying members; 90% probability with 95% confidence. (c) The effects of temperature on allowable stresses used for design in an essential component or structure must be considered where thermal effects are significant under normal operating conditions. (d) The design of structure must minimise the probability of catastrophic fatigue failure, particularly at points of stress concentration. (e) Design values greater than the guaranteed minimum’s required by this paragraph may be used where only guaranteed minimum values are normally allowed if a “premium selection” of the material is made in which J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 19 of 30 JC/DOC/325 ISSUE 1.4 a specimen of each individual item is tested before use to determine that the actual strength properties of the particular item will equal or exceed those used in design. The material strength properties are based on approved specifications. The existing structure and fasteners have been assessed in JC/TR/54. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.619 Special Factors: The factor of safety prescribed in CS 23.303 must be multiplied by the highest pertinent special factors of safety prescribed in CS 23.621 to 23.625 for each part of the structure whose strength is – (1) Uncertain; (2) Likely to deteriorate in service before normal replacement; or (3) Subject to appreciable variability because of uncertainties in manufacturing processes or inspection methods. Not affected by this change. A static test has been carried out to ensure the structural integrity of the XCAM installation under flight and ground induced loads. See JC/TR/54 as defined in section 11. CS23.621 Casting Factors: (a) General. The factors, tests and inspections specified in subparagraphs (b) to (d) must be applied in addition to those necessary to establish foundry quality control. The inspections must meet approved specifications. Subparagraphs (c) and (d) apply to any structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads.(c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the aeroplane or result in serious injury to occupants, the following apply: Each critical casting must either (i) Have a casting factor of not less than 1∙25 and receive 100% inspection by visual, radiographic and either magnetic particle, penetrant or other approved equivalent non- destructive inspection method or (ii) Have a casting factor of not less than 2 ∙0 and receive 100% visual inspection and 100% approved non-destructive inspection. When an approved quality control procedure is established and an acceptable statistical analysis supports reduction, non-destructive inspection may be reduced from 100%, and applied on a sampling basis. No affected by this change. Not applicable as the part is not cast. See ECN(s)/Drawing(s) as defined in section 11. CS23.623 Bearing Factors: (a) Each part that has clearance (free fit) and that is subject to pounding or vibration, must have a bearing factor large enough to provide for the effects of normal relative motion. (b) For control surface hinges and control system joints, compliance with the factors prescribed in CS 23.657 and 23.693 respectively, meets paragraph (a) . A static test has been carried out to ensure the structural integrity of the XCAM installation under flight and ground induced loads. See JC/TR/54 as defined in section 11. CS23.625 Fitting Factors: For each fitting (a part or terminal used to join one structural member to another), the following applies: (a) For each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1 ∙15 must be applied to each part of – (1) The fitting; (2) The means of attachment; and (3) The bearing on the joined members. (b) No fitting factor need be used for joint designs based on comprehensive test data (such as continuous joints in metal plating, welded joints and scarf joints in wood). (c) For each integral fitting, the part must be treated as a fitting up to the point at which the section properties become typical of the member. (d) For each seat, berth, safety belt and harness, its attachment to the structure must be shown, by analysis, tests, or both, to be able to withstand the inertia forces prescribed in CS 23.561 multiplied by a fitting factor of 1 ∙33. J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 20 of 30 JC/DOC/325 ISSUE 1.4 A static test has been carried out to ensure the structural integrity of the XCAM installation under flight and ground induced loads. See JC/TR/54 as defined in section 11. CS23.627 Fatigue Strength: The structure must be designed, as far as practicable, to avoid points of stress concentration where variable stresses above the fatigue limit are likely to occur in normal service. Not affected by this change. The installation of the XCAM via the clamping arrangement does not introduce any new stress concentrations. See ECN(s)/Drawing(s) and JC/TR/54 as defined in section 11. CS23.629 Flutter: (a) It must be shown by the methods of (b) and either (c) or (d) , that the aeroplane is free from flutter, control reversal and divergence for any condition of operation within the limit V-n envelope and at all speeds up to the speed specified for the selected method. In addition – (1) Adequate tolerances must be established for quantities which affect flutter; including speed, damping, mass balance and control system stiffness; and (2) The natural frequencies of main structural components must be determined by vibration tests or other approved methods. (b) Flight flutter tests must be made to show that the aeroplane is free from flutter, control reversal and divergence and to show by these tests that – (1) Proper and adequate attempts to induce flutter have been made within the speed range up to VD; (2) The vibratory response of the structure during the test indicates freedom from flutter; (3) A proper margin of damping exists at VD; and (4) There is no large and rapid reduction in damping as VD is approached. (c) Any rational analysis used to predict freedom from flutter, control reversal and divergence must cover all speeds up to 1·2 VD. (d) Compliance with the rigidity and mass balance criteria (pages 4-12), in Airframe and Equipment Engineering Report No. 45 (as corrected) “Simplified Flutter Prevention Criteria” (published by the Federal Aviation Administration) may be accomplished to show that the aeroplane is free from flutter, control reversal, or divergence if – (1) VD/MD for the aeroplane is less than 482 km/h (260 knots) (EAS) and less than Mach 0·5; (2) The wing and aileron flutter prevention criteria, as represented by the wing torsional stiffness and aileron balance criteria, are limited to use to aeroplanes without large mass concentrations (such as engines, floats, or fuel tanks in outer wing panels) along the wing span; and (3) The aeroplane – (i) Does not have a T-tail or other unconventional tail configurations; (ii) Does not have unusual mass distributions or other unconventional design features that affect the applicability of the criteria; and (iii) Has fixed-fin and fixed stabiliser surfaces. Not affected by this change. The airframe, flying control surfaces and systems stiffness have not been changed by the introduction of the XCAM installation and clamping arrangement. See JC/TR/54 as defined in section 11. CS23.787 Baggage and Cargo Compartments: (a) Each baggage and cargo compartment must – (1) Be designed for its placarded maximum weight of contents and for the critical load distributions at the appropriate maximum load factors corresponding to the flight and ground load conditions of CS-23. (2) Have means to prevent the contents of any compartment from becoming a hazard by shifting, and to protect any controls, wiring, lines, equipment, or accessories whose damage or failure would affect safe operations. (3) Have a means to protect occupants from injury by the contents of any compartment, located aft of the occupants and separated by structure, when the ultimate forward inertia load factor is 9g and assuming the maximum allowed baggage or cargo weight for the compartment. (b) Aeroplanes that provide for baggage or cargo to be carried in the same compartment as passengers must have a means to protect the occupants from injury when the baggage or cargo is subjected to the inertia loads resulting from the ultimate static load factors of CS 23.561 (b) (3), assuming J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 21 of 30 JC/DOC/325 ISSUE 1.4 the maximum allowed baggage or cargo weight for the compartment. (c) For aeroplanes that are used only for the carriage of cargo, the flight crew emergency exits must meet the requirements of CS 23.807 under any baggage or cargo loading conditions. The PC/Tablet and Battery Pack are stowed for Taxi, Take Off and Landing in an already approved stowage location. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.791 Passenger Information Signs: For those aeroplanes in which the flight crew members can not observe the other occupants seats or in which the crew compartment is separated from the passenger compartment, there must be at least one illuminated sign (using either letters or symbols) notifying all passengers when safety belts must be fastened. Signs that notify when seat belts should be fastened must – (a) When illuminated, be legible to each person seated in the passenger compartment under all probable lighting conditions; and (b) Be installed so that a flight-crew member can, when seated at their station, turn the illumination on and off. Not affected by this change. The aircraft can only be operated in the Aerial Work Regime when the XCAM is installed. See JC/FMS/4584 as defined in section 11. CS23.803 Emergency Evacuation: (a) For commuter category aeroplanes, an evacuation demonstration must be conducted utilising the maximum number of occupants for which certification is desired. The demonstration must be conducted under simulated night conditions using only the emergency exits on the most critical side of the aeroplane. The participants must be representative of average airline passengers with no prior practice or rehearsal for the demonstration. Evacuation must be completed within 90 seconds. (b) In addition, when certification to the emergency exit provisions of CS 23.807(d)(4) is requested, only the emergency lighting system required by CS 23.812 may be used to provide cabin interior illumination during the evacuation demonstration required in sub-paragraph (a) . Not affected by this change. The aircraft can only be operated in the Aerial Work Regime when the XCAM is installed. See JC/FMS/4584 as defined in section 11. CS23.805 Flight Crew Emergency Exits: For aeroplanes where the proximity of the passenger emergency exits to the flightcrew area does not offer a convenient and readily accessible means of evacuation for the flightcrew, the following apply: (a) There must be either one emergency exit on each side of the aeroplane, or a top hatch emergency exit, in the flightcrew area; (b) Each emergency exit must be located to allow rapid evacuation of the crew and have a size and shape of at least a 48-by 51 cm (19- by 20-in) unobstructed rectangular opening; and (c) For each emergency exit that is not less than 1·8 metres (6 ft) from the ground, an assisting means must be provided. The assisting means may be a rope or any other means demonstrated to be suitable for the purpose. If the assisting means is a rope or an approved device equivalent to a rope, it must be- (1) Attached to the fuselage structure at or above the top of the emergency exit opening or, for a device at a pilot's emergency exit window, at another approved location if the stowed device, or its attachment, would reduce the pilot's view; and (2) Able (with its attachment) to withstand a 1779 N (400 lbf) static load. Not affected by this change. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.807 Emergency Exits: a) Number and location. Emergency exits must be located to allow escape without crowding in any probable crash attitude. The aeroplane must have at least the following emergency exits: (1) For all aeroplanes with a seating capacity of two or J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 22 of 30 JC/DOC/325 ISSUE 1.4 more, excluding aeroplanes with canopies, at least one emergency exit on the opposite side of the cabin from the main door specified in CS 23.783. Not affected by this change. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.813 Emergency Exit Access: (a) For commuter category aeroplanes, access to window-type emergency exits may not be obstructed by seats or seat backs. (b) In addition, when certification to the emergency exit provisions of CS 23.807(d)(4) is requested, the following emergency exit access must be provided: (1) The passageway leading from the aisle to the passenger entry door must be unobstructed and at least 51 cm (20 in) wide. (2) There must be enough space next to the passenger entry door to allow assistance in evacuation of passengers without reducing the unobstructed width of the passageway below 51 cm (20 in). (3) If it is necessary to pass through a passageway between passenger compartments to reach a required emergency exit from any seat in the passenger cabin, the passageway must be unobstructed; however, curtains may be used if they allow free entry through the passageway. (4) No door may be installed in any partition between passenger compartments unless that door has a means to latch it in the open position. The latching means must be able to withstand the loads imposed upon it by the door when the door is subjected to the inertia loads resulting from the ultimate static load factors prescribed in CS 23.561(b)(2). (5) If it is necessary to pass through a door-way separating the passenger cabin from other areas to reach a required emergency exit from any passenger seat, the door must have a means to latch it in the open position. The latching means must be able to withstand the loads imposed upon it by the door when the door is subjected to the inertia loads resulting from the ultimate static load factors prescribed in CS 23.561(b)(2). Not affected by this change. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.815 Width of Aisle: a) Except as provided in sub-paragraph (b), for commuter category aeroplanes, the width of the main passenger aisle at any point between seats must equal or exceed the values in the following table: Not affected by this change. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS23.1501 Operating Limitations and Information General: (a) Each operating limitation specified in CS 23.1505 to 23.1527 and other limitations and information necessary for safe operation must be established. (b) The operating limitations and other information necessary for safe operation must be made available to the crew members as prescribed in CS 23.1541 to 23.1589. The introduction of the XCAM does not change the operating limitation as defined in JC/FMS/4584. The aircraft can only be operated in the Aerial Work Regime when the XCAM in installed. See JC/FMS/4584 as defined in section 11. CS 23.1505 Airspeed Limitations: (a) The never-exceed speed VNE must be established so that it is – (1) Not less than 0·9 times the minimum value of VD allowed under CS 23.335; and (2) Not more than the lesser of – (i) 0·9 VD established under CS 23.335; or (ii) 0·9 times the maximum speed shown under CS 23.251. (b) The maximum structural cruising speed VNO must be established so that it is – (1) Not less than the minimum value of VC allowed under CS 23.335; and (2) Not more than the lesser of – (i) VC established under CS 23.335; or (ii) 0·89 VNE established under sub-paragraph (a) . (c) Sub-paragraphs (a) and (b) do not apply to turbine aeroplanes or to aeroplanes for which a design diving speed VD/MD is established J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 23 of 30 JC/DOC/325 ISSUE 1.4 under CS 23.335 (b) (4). For those aeroplanes, a maximum operating limit speed (VMO/MMO airspeed or Mach number, whichever is critical at a particular altitude) must be established as a speed that may not be deliberately exceeded in any regime of flight (climb, cruise, or descent) unless a higher speed is authorised for flight test or pilot training operations. VMO/MMO must be established so that it is not greater than the design cruising speed VC/MC and so that it is sufficiently below VD/MD and the maximum speed shown under CS 23.251 to make it highly improbable that the latter speeds will be inadvertently exceeded in operations. The speed margin between VMO/MMOand VD/MD or the maximum speed shown under CS 23.251 may not be less than the speed margin established between VC/MC and VD/MD under CS 23.335(b), or the speed margin found necessary in the flight tests conducted under CS 23.253. Not affected by this change. See JC/FMS/4584 as defined in section 11. CS 23.1507 Manoeuvring Speed: The maximum operating manoeuvring speed, VO, must be established as an operating limitation. VO is a selected speed that is not greater than VS √n established in CS 23.335(c). Not affected by this change. See JC/FMS/4584 as defined in section 11. CS23.1519 Weight and Centre of Gravity: The weight and centre of gravity ranges determined under CS 23.23 must be established as operating limitations. There is a change to the weight and balance by the introduction of XCAM pod. See section 7.0. CS23.1529 Instructions for Continued Airworthiness: Instructions for continued airworthiness in accordance with Appendix G must be prepared. The installation is to be inspected IAW operator’s and manufacturers requirements as defined in the AMM/CMM and SRM. Co-ordination of these matters is the responsibility of the operator. See section 10. CS 23.1541 Markings and Placards General: (a) The aeroplane must contain – (1) The markings and placards specified in CS 23.1545 to 23.1567; and (2) Any additional information, instrument markings and placards required for the safe operation if it has unusual design, operating, or handling characteristics. (b) Each marking and placard prescribed in sub-paragraph (a) – (1) Must be displayed in a conspicuous place; and (2) May not be easily erased, disfigured or obscured. (c) For aeroplanes which are to be certificated in more than one category – (1) One category upon which the placards and markings are to be based must be selected for the aeroplane; and (2) The placards and marking information for all categories in which the aeroplane is to be certificated must be furnished in the Aeroplane Flight Manual. Not affected by this change. See ECN(s)/Drawing(s) and JC/FMS/4584 as defined in section 11. CS 23.1581 Aeroplane Flight Manual General: (a) An Aeroplane Flight Manual must be submitted to the Agency and it must contain the following: (1) Information required by CS 23.1583 to 23.1589. (2) Other information that is necessary for safe operation because of design, operating or handling characteristics. (3) Further information necessary to comply with the relevant operating rules. (b) Approved information (1) Except as provided in sub-paragraph (b)(2), each part of the Aeroplane Flight Manual containing information prescribed in CS 23.1583 to 23.1589 must be approved, segregated, identified and clearly distinguished from each unapproved part of that Aeroplane Flight Manual. (2) The requirements of sub-paragraph (b) (1) J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 24 of 30 JC/DOC/325 ISSUE 1.4 do not apply to reciprocating engine powered aeroplanes of 2 722 kg (6 000 lb) or less maximum weight, if the following is met: (i) Each part of the Aeroplane Flight Manual containing information prescribed in CS 23.1583 must be limited to such information and must be approved, identified and clearly distinguished from each other part of the Aeroplane Flight Manual. (ii) The information prescribed in CS 23.1585 to 23.1589 must be determined in accordance with the applicable requirements of CS-23 and presented in its entirety in a manner acceptable to the Agency. (c) The units used in the Aeroplane Flight Manual must be the same as those marked on the appropriate instruments and placards. (d) All Aeroplane Flight Manual operational airspeeds must, unless otherwise specified, be presented as indicated Airspeeds. (e) Provisions must be made for stowing the Aeroplane Flight Manual in a suitable fixed container which is readily accessible to the pilot. (f) Revisions and/or Amendments. Each Aeroplane Flight Manual must contain a means for recording the incorporation The delta change in performance of the aircraft has been assessed in JC/TR/54. An AFMS has been raised JC/FMS/4584 to define its installation and operational limitations. See JC/TR/54 and JC/FMS/4584 as defined in section 11. CS 23.1583 Operating Limitations: The Aeroplane Flight Manual must contain operating limitations determined under CS-23, including the following: (a) Airspeed limitations (1) Information necessary for the marking of the airspeed limits on the indicator as required in CS 23.1545, and the significance of each of those limits and of the colour coding used on the indicator. (2) The speeds VMC, Vo, VLE and VLO and their significance. (3) In addition, for turbine powered commuter category aeroplanes – (i) The maximum operating limit speed, VMO/MMO and a statement that this speed must not be deliberately exceeded in any regime of flight (climb, cruise or descent) unless a higher speed is authorised for flight test or pilot training; (ii) If an airspeed limitation is based upon compressibility effects, a statement to this effect and information as to any symptoms, the probable behaviour of the aeroplane and the recommended recovery procedures; and (iii) The airspeed limits must be shown in terms of VMO/MMO instead of VNO and VNE. (b) Powerplant limitations (1) Limitations required by CS 23.1521. (2) Explanation of the limitations, when appropriate. (3) Information necessary for marking the instruments required by CS 23.1549 to 23.1553. (c) Weight (1) The maximum weight; and (2) The maximum landing weight, if the design landing weight selected by the applicant is less than the maximum weight. (3) For normal, utility and aerobatic category reciprocating engine-powered aeroplanes of more than 2 722 kg (6 000 lb) maximum weight and for turbine engine powered aeroplanes in the normal, utility and aerobatic category, performance operating limitations as follows: (i) The maximum take-off weight for each aerodrome altitude and ambient temperature within the range selected by the applicant at which the aeroplane complies with the climb requirements of CS 23.63 (c) (1). (ii) The maximum landing weight for each aerodrome altitude and ambient temperature within the range selected by the applicant at which the aeroplane complies with the climb requirements of CS 23.63 (c) (2). (4) For commuter category aeroplanes, the maximum take-off weight for each aerodrome altitude and ambient temperature within the range selected by the applicant at which – (i) The aeroplane complies with the climb requirements of CS 23.63 (d) (1); and (ii) The accelerate-stop distance determined under CS 23.55 is equal to the available runway length plus the length of any stop way, if utilised; and either, (iii) The take-off distance determined under CS 23.59 (a) is equal to the available runway length; or (iv) At the option of the applicant, the take-off distance determined under CS 23.59 (a) is equal to the available runway length plus the length of any clearway and the take-off run determined under CS 23.59 (b) is equal to the available runway length. (5) For commuter category aeroplanes, the maximum landing weight for each aerodrome altitude within the range selected by the applicant at which – (i) The aeroplane complies with the climb requirements of CS 23.63(d)(2) for ambient temperatures within the range selected by the applicant. (ii) The landing distance determined under CS 23.75 for standard temperatures is equal to the available runway J & C Engineering Services Ltd ENGINEERING REPORT Report No. Issue JC/ER/4584 1.5 Sheet 25 of 30 JC/DOC/325 ISSUE 1.4 length; and (6) The maximum zero wing fuel weight where relevant as established in accordance with CS 23.343. (d) Centre of gravity. The established centre of gravity limits. (e) Manoeuvres. The following authorised manoeuvres, appropriate airspeed limitations, and unauthorised manoeuvres, as prescribed in this section. (1) Normal category aeroplanes. No aerobatic manoeuvres, including spins, are authorised. (2) Utility category aeroplanes. A list of authorised manoeuvres demonstrated in the type flight tests, together with recommended entry speeds and any other associated limitations. No other manoeuvre is authorised. (3) Aerobatic category aeroplanes. A list of approved flight manoeuvres demonstrated in the type flight tests, together with reco