Idition it must be shown that the airplane

safely controllable and that a pilot can erform all the maneuvers and operations ecessary to effect a safe landing following ny probable electric trim tab runaway which might be reasonably expected in servce allowing for appropriate time delay after ilot recognition of the runaway. This lemonstration must be conducted at the ritical airplane weights and center of gravty positions.

INSTRUMENTS: INSTALLATION

12. Arrangement and visibility. Each intrument must meet the requirements of FAR 23.1321 and in addition

(a) Each flight, navigation, and powerlant instrument for use by any pilot must e plainly visible to him from his station with the minimum practicable deviation rom his normal position and line of vision when he is looking forward along the flight ath.

(b) The flight instruments required by AR 23.1303 and by the applicable operating ules must be grouped on the instrument panel and centered as nearly as practicable about the vertical plane of each pilot's forward vision. In addition

(1) The instrument that most effectively Indicates the attitude must be on the panel En the top center position;

(2) The instrument that most effectively ndicates airspeed must be adjacent to and directly to the left of the instrument in the top center position;

(3) The instrument that most effectively indicates altitude must be adjacent to and directly to the right of the instrument in the top center position; and

(4) The instrument that most effectively indicates direction of flight must be adjacent to and directly below the instrument in the Top center position.

13. Airspeed indicating system. Each airspeed indicating system must meet the requirements of FAR 23.1323 and in addition— (a) Airspeed indicating instruments must De of an approved type and must be callDrated to indicate true airspeed at sea level n the standard atmosphere with a minimum practicable instrument calibration error Then the corresponding pilot and static presures are supplied to the instruments.

(b) The airspeed indicating system must De calibrated to determine the system error, i.e., the relation between IAS and CAS, in flight and during the accelerate takeoff ground run. The ground run calibration must be obtained between 0.8 of the minimum value of V1 and 1.2 times the maximum value of V1, considering the approved ranges of altitude and weight. The ground run caliration will be determined assuming an engine failure at the minimum value of V1.

(c) The airspeed error of the installation excluding the instrument calibration error, must not exceed 3 percent or 5 knots whichever is greater, throughout the speed range

from Vмo to 1.3S1 with flaps retracted and from1.3V So to VFE with flaps in the landing position.

(d) Information showing the relationship between IAS and CAS must be shown in the Airplane Flight Manual.

14. Static air vent system. The static air vent system must meet the requirements of FAR 23.1325. The altimeter system calibration must be determined and shown in the Airplane Flight Manual.

OPERATING LIMITATIONS AND INFORMATION

15. Maximum operating limit speed Vμо/ Mмo. Instead of establishing operating limitations based on VNE and VNo, the applicant must establish a maximum operating limit speed VMO/Mмo in accordance with the following:

(a) The maximum operating limit speed must not exceed the design cruising speed Vc and must be sufficiently below VD/MD or VDF/MDF to make it highly improbable that the latter speeds will be inadvertently exceeded in flight.

(b) The speed Vмo must not exceed 0.8VD/MD or 0.8VDF/MDF unless flight demonstrations involving upsets as specified by the Administrator indicates a lower speed margin will not result in speeds exceeding VD/MD or VDF. Atmospheric variations, horizontal gusts, and equipment errors, and airframe production variations will be taken into account.

16. Minimum flight crew. In addition to meeting the requirements of FAR 23.1523, the applicant must establish the minimum number and type of qualified flight crew personnel sufficient for safe operation of the airplane considering

(a) Each kind of operation for which the applicant desires approval;

(b) The workload on each crewmember considering the following:

(1) Flight path control. (2) Collision avoidance. (3) Navigation.

(4) Communications.

(5) Operation and monitoring of all essential aircraft systems.

(6) Command decisions; and

(c) The accessibility and ease of operation of necessary controls by the appropriate crewmember during all normal and emergency operations when at his flight station.

17. Airspeed indicator. The airspeed indicator must meet the requirements of FAR 23.1545 except that, the airspeed notations and markings in terms of Vro and VNE must be replaced by the Vо/MMо notations. The airspeed indicator markings must be easily read and understood by the pilot. A placard adjacent to the airspeed indicator is an acceptable means of showing compliance with the requirements of FAR 23.1545 (c). AIRPLANE FLIGHT MANUAL

18. General. The Airplane Flight Manual must be prepared in accordance with the requirements of FARS 23.1583 and 23.1587, and

in addition the operating limitations and performance information set forth in sections 19 and 20 must be included.

19. Operating limitations. The Airplane Flight Manual must include the following limitations—

(a) Airspeed limitations. (1) The maximum operating limit speed VMо/Mмо and a statement that this speed limit may not be deliberately exceeded in any regime of flight (climb, cruise, or descent) unless a higher speed is authorized for flight test or pilot training;

(2) If an airspeed limitation is based upon compressibility effects, a statement to this effect and information as to any symptoms, the probable behavior of the airplane, and the recommended recovery procedures; and

(3) The airspeed limits, shown in terms of VMO/MMо instead of VNO and VNE.

(b) Takeoff weight limitations. The maximum takeoff weight for each airport elevation, ambient temperature, and available takeoff runway length within the range selected by the applicant. This weight may not exceed the weight at which:

(1) The all-engine operating takeoff distance determined in accordance with section 5(d) or the accelerate-stop distance determined in accordance with section 5(c), which ever is greater, is equal to the available runway length;

(2) The airplane complies with the oneengine-inoperative takeoff requirements specified in § 5(e); and

(3) The airplane complies with the oneengine-inoperative en route climb requirements specified in § 6(b), assuming that a standard temperature lapse rate exists from the airport elevation to the altitude of 5,000 feet, except that the weight may not exceed that corresponding to a temperature of 41° F. at 5,000 feet.

20. Performance information. The Airplane Flight Manual must contain the performance information determined in accordance with the provisions of the performance requirements of this regulation. The information must include the following:

(a) Sufficient information so that the takeoff weight limits specified in § 19 (b) can be determined for all temperatures and altitudes within the operation limitations selected by the applicant.

(b) The conditions under which the performance information was obtained, including the airspeed at the 50-foot height used to determine landing distances.

(c) The performance information (determined by extrapolation and computed for the range of weights between the maximum landing and takeoff weights) for

(1) Climb in the landing configuration; and

(2) Landing distance.

(d) Procedure established under section 4 of this regulation related to the limitations and information required by this

section in the form of guidance materia including any relevant limitations 0 information.

(e) An explanation of significant or un usual flight or ground handling character istics of the airplane.

(f) Airspeeds, as indicated airspeeds corresponding to those determined for take off in accordance with section 5(b).

21. Maximum operating altitudes. Th maximum operating altitude to which opera tion is permitted, as limited by flight, struc tural, powerplant, functional, or equipmen characteristics, must be specified in the Air plane Flight Manual.

22. Stowage provision for Airplane Fligh Manual. Provision must be made for stow ing the Airplane Flight Manual in a suitabl fixed container which is readily accessibl to the pilot.

23. Operating procedures. Procedures fo restarting turbine engines in flight (includin the effects of altitude) must be set forth i the Airplane Flight Manual.

AIRFRAME REQUIREMENTS

FLIGHT LOADS

24. Engine torque. (a) Each turbopro peller engine mount and its supporting struc ture must be designed for the torque effect of

(1) The conditions set forth in FAR 23. 361(a).

(2) The limit engine torque correspond ing to takeoff power and propeller speed multiplied by a factor accounting for propeller control system malfunction, including quick feathering action, simultaneously with 1g level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used

(b) The limit torque is obtained by mul tiplying the mean torque by a factor of 1.25

25. Turbine engine gyroscopic loads. Each turbopropeller engine mount and its support ing structure must be designed for the gyro scopic loads that result, with the engines a maximum continuous r.p.m., under either(a) The conditions prescribed in FAR 23.351 and 23.423; or

(b) All possible combinations of the fol lowing:

(1) A yaw velocity of 2.5 radius per second (2) A pitch velocity of 1.0 radians pe second.

(3) A normal load factor of 2.5

(4) Maximum continuous thrust. 26. Unsymmetrical loads due to engine fail ure. (a) Turbopropeller powered airplane must be designed for the unsymmetrica loads resulting from the failure of the critical engine including the following conditions in combination with a single malfunc tion of the propeller drag limiting system considering the probable pilot correctiv action on the flight controls.

(1) At speeds between Vme and VD, th

pads resulting from power failure because f fuel flow interruption are considered to be limit loads.

(2) At speeds between Vmo and Vc, the loads resulting from the disconnection of the engine compressor from the turbine or from loss of the turbine blades are considered to be ultimate loads.

(3) The time history of the thrust decay and drag buildup occurring as a result of the prescribed engine failures must be substantiated by test or other data applicable to the particular engine-propeller combination.

(4) The timing and magnitude of the probable pilot corective action must be conservatively estimated, considering the characteristics of the particular engine-propellerairplane combination.

(b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than two seconds after the engine failure. The magnitude of the corrective action may be based on the control forces specified in FAR 23.397 except that lower forces may be assumed where it is shown by analysis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions.

GROUND LOADS

27. Dual wheel landing gear units. Each dual wheel landing gear unit and its supporting structure must be shown to comply with the following:

(a) Pivoting. The airplane must be assumed to pivot about one side of the main gear with the brakes on that side locked. The limit vertical load factor must be 1.0 and the coefficient of friction 0.8. This condition need apply only to the main gear and its supporting structure.

(b) Unequal tire inflation. A 60-40 percent distribution of the loads established in accordance with FAR 23.471 through FAR 23.483 must be applied to the dual wheels. (c) Flat tire. (1) Sixty percent of the loads specified in FAR 23.471 through FAR 23.483 must be applied to either wheel in a

unit.

(2) Sixty percent of the limit drag and side loads and 100 percent of the limit vertical load established in accordance with FARS 23.493 and 23.485 must be applied to either wheel in a unit except that the vertical load need not exceed the maximum vertical load in paragraph (c) (1) of this section.

FATIGUE EVALUATION

28. Fatigue evaluation of wing and associated structure. Unless it is shown that the Structure, operating stress levels, materials, and expected use are comparable from a fatigue standpoint to a similar design which has had substantial satisfactory service experience, the strength, detail design, and the fabrication of those parts of the wing, wing

36-030-70- -6

carrythrough, and attaching structure whose failure would be catastrophic must be evaluated under either

(a) A fatigue strength investigation in which the structure is shown by analysis, tests, or both 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, 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 Vo. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise considered.

DESIGN AND CONSTRUCTION

29. Flutter. For Multiengine turbopropeller powered airplanes, a dynamic evaluation must be made and must include

(a) The significant elastic, inertia, and aerodynamic forces associated with the rotations and displacements of the plane of the propeller; and

(b) Engine-propeller-nacelle stiffness and damping variations appropriate to the particular configuration.

LANDING GEAR

30. Flap operated landing gear warning device. Airplanes having retractable landing gear and wing flaps must be equipped with a warning device that functions continuously when the wing flaps are extended to a flap position that activates the warning device to give adequate warning before landing, using normal landing procedures, if the landing gear is not fully extended and locked. There may not be a manual shut off for this warning device. The flap position sensing unit may be installed at any suitable location. The system for this device may use any part of the system (including the aural warning device) provided for other landing gear warning devices.

PERSONNEL AND CARGO ACCOMMODATIONS

31. Cargo and baggage compartments. Cargo and baggage compartments must be designed to meet the requirements of FAR 23.787 (a) and (b), and in addition means must be provided to protect passengers from injury by the contents of any cargo or baggage compartment when the ultimate forward inertia force is 9g.

32. Doors and exits. The airplane must meet the requirements of FAR 23.783 and FAR 23.807 (a)(3), (b), and (c), and in addition:

(a) There must be a means to lock and safeguard each external door and exit against opening in flight either inadvertently by

persons, or as a result of mechanical failure. Each external door must be operable from both the inside and the outside.

(b) There must be means for direct visual inspection of the locking mechanism by crewmembers to determine whether external doors and exits, for which the initial opening movement is outward, are fully locked. In addition, there must be a visual means to signal to crewmembers when normally used external doors are closed and fully locked.

(c) The passenger entrance door must qualify as a floor level emergency exit. Each additional required emergency exit except floor level exits must be located over the wing or must be provided with acceptable means to assist the occupants in descending to the ground. In addition to the passenger entrance door:

(1) For a total seating capacity of 15 or less, an emergency exit as defined in FAR 23.807(b) is required on each side of the cabin.

(2) For a total seating capacity of 16 through 23, three emergency exits as defined in 23.807(b) are required with one on the same side as the door and two on the side opposite the door.

(d) An evacuation demonstration must be conducted utilizing the maximum number of occupants for which certification is desired. It must be conducted under simulated night conditions utilizing only the emergency exits on the most critical side of the aircraft. 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.

(e) Each emergency exit must be marked with the word "Exit" by a sign which has white letters 1 inch high on a red background 2 inches high, be self-illuminated or independently internally electrically illuminated, and have a minimum luminescence (brightness) of at least 160 microlamberts. The colors may be reversed if the passenger compartment illumination is essentially the

frame must be connected to it through ligh ning arrestors unless a lightning strike the insulated part

(a) Is improbable because of shielding other parts; or

(b) Is not hazardous.

34. Ice protection. If certification with i protection provisions is desired, complian with the following requirements must shown:

(a) The recommended procedures for use of the ice protection equipment must set forth in the Airplane Flight Manual,

(b) An analysis must be performed to e tablish, on the basis of the airplane's oper tional needs, the adequacy of the ice pr tection system for the various componen of the airplane. In addition, tests of the i protection system must be conducted demonstrate that the airplane is capable operating safely in continuous maximu and intermittent maximum icing condition as described in FAR 25, Appendix C.

(c) Compliance with all or portions of th section may be accomplished by referend where applicable because of similarity the designs, to analysis and tests performe by the applicant for a type certificated mode

35. Maintenance information. The appl cant must make available to the owner a the time of delivery of the airplane the in formation he considers essential for th proper maintenance of the airplane. Tha information must include the following:

(a) Description of systems, including elec trical, hydraulic, and fuel controls.

(b) Lubrication instructions setting forth the frequency and the lubricants and fluids which are to be used in the various systems

(c) Pressures and electrical loads appli cable to the various systems.

(d) Tolerances and adjustments necessar for proper functioning.

(e) Methods of leveling, raising, an towing.

(f) Methods of balancing control surfaces (g) Identification of primary and second ary structures.

(h) Frequency and extent of inspection necessary to the proper operation of th airplane.

(i) Special repair methods applicable t the airplane.

(j) Special inspection techniques, includ ing those that require X-ray, ultrasonic, an magnetic particle inspection.

(k) List of special tools.

PROPULSION

GENERAL

36. Vibration characteristics. For turbopropeller powered airplanes, the engine installation must not result in vibration char acteristics of the engine exceeding thos established during the type certification o the engine.

37. In-flight restarting of engine. If the engine on turbopropeller powered airplane

annot be restarted at the maximum cruise altitude, a determination must be made of the altitude below which restarts can be consistently accomplished. Restart information must be provided in the Airplane Flight Manual.

38. Engines (a) For turbopropeller powered airplanes. The engine installation must comply with the following requirements:

(1) Engine isolation. The powerplants must be arranged and isolated from each other to allow operation, in at least one configuration, so that the failure or malfunction of any engine, or of any system that can affect the engine, will not

(i) Prevent the continued safe operation of the remaining engines; or

(ii) Require immediate action by any crewmember for continued safe operation.

(2) Control of engine rotation. There must be a means to individually stop and restart the rotation of any engine in flight except that engine rotation need not be stopped if continued rotation could not jeopardize the safety of the airplane. Each component of the stopping and restarting system on the engine side of the firewall, and that might be exposed to fire, must be at least fire resistant. If hydraulic propeller feathering systems are used for this purpose, the feathering lines must be at least fire resistant under the operating conditions that may be expected to exist during feathering.

(3) Engine speed and gas temperature control devices. The powerplant systems associated with engine control devices, systems, and instrumentation must provide reasonable assurance that those engine operating limitations that adversely affect turbine rotor structural integrity will not be exceeded in

service.

(b) For reciprocating-engine powered airplanes. To provide engine isolation, the powerplants must be arranged and isolated from each other to allow operation, in at least one configuration, so that the failure or malfunction of any engine, or of any system that can affect that engine, will not

(1) Prevent the continued safe operation of the remaining engines; or

(2) Require immediate action by any crewmember for continued safe operation.

39. Turbopropeller reversing systems. (a) Turbopropeller reversing systems intended for ground operation must be designed so that no single failure or malfunction of the system will result in unwanted reverse thrust under any expected operating condition. Failure of structural elements need not De considered if the probability of this kind of failure is extremely remote.

(b) Turbopropeller reversing systems intended for in-flight use must be designed so that no unsafe condition will result during normal operation of the system, or from any failure (or reasonably likely combination of failures) of the reversing system, under any anticipated condition of operation of the airplane. Failure of structural elements need

not be considered if the probability of this kind of failure is extremely remote.

(c) Compliance with this section may be shown by failure analysis, testing, or both for propeller systems that allow propeller blades to move from the flight low-pitch position to a position that is substantially less than that at the normal flight low-pitch stop position. The analysis may include or be supported by the analysis made to show compliance with the type certification of the propeller and associated installation components. Credit will be given for pertinent analysis and testing completed by the engine and propeller manufacturers.

40. Turbopropeller drag-limiting systems. Turbopropeller drag-limiting systems must be designed so that no single failure or malfunction of any of the systems during normal or emergency operation results in propeller drag in excess of that for which the airplane was designed. Failure of structural elements of the drag-limiting systems need not be considered if the probability of this kind of failure is extremely remote.

41. Turbine engine powerplant operating characteristics. For turbopropeller powered airplanes, the turbine engine powerplant operating characteristics must be investigated in flight to determine that no adverse characteristics (such as stall, surge, or flameout) are present to a hazardous degree, during normal and emergency operation within the range of operating limitations of the airplane and of the engine.

42. Fuel flow. (a) For turbopropeller powered airplanes

(1) The fuel system must provide for continuous supply of fuel to the engines for normal operation without interruption due to depletion of fuel in any tank other than the main tank; and

(2) The fuel flow rate for turbopropeller engine fuel pump systems must not be less than 125 percent of the fuel flow required to develop the standard sea level atmospheric conditions takeoff power selected and included as an operating limitation in the Airplane Flight Manual.

(b) For reciprocating engine powered airplanes, it is acceptable for the fuel flow rate for each pump system (main and reserve supply) to be 125 percent of the takeoff fuel consumption of the engine.

FUEL SYSTEM COMPONENTS

43. Fuel pumps. For turbopropeller powered airplanes, a reliable and independent power source must be provided for each pump used with turbine engines which do not have provisions for mechanically driving the main pumps. It must be demonstrated that the pump installations provide a reliability and durability equivalent to that provided by FAR 23.991(a).

44. Fuel strainer or filter. For turbopropeller powered airplanes, the following apply:

(a) There must be a fuel strainer or filter