The reliability of any spring device used in the control system must be established by tests simulating service conditions unless failure of the spring will not cause flutter or unsafe flight characteristics.

$23.689 Cable systems.

(a) Each cable, cable fitting, turnuckle, splice, and pulley used must meet pproved specifications. In addition-

(1) No cable smaller than 8 inch dimeter may be used in primary control ystems;

(2) Each cable system must be deigned so that there will be no hazardous hange in cable tension throughout the ange of travel under operating condilons and temperature variations; and

(3) There must be means for visual Inspection at each fairlead, pulley, terninal, and turnbuckle.

(b) Each kind and size of pulley must orrespond to the cable with which it is ised, as specified in the pulley specificaion. Each pulley must have closely itted guards to prevent the cables from eing misplaced or fouled, even when lack. Each pulley must lie in the plane assing through the cable so that the able does not rub against the pulley lange.

(c) Fairleads must be installed so that they do not cause a change in cable direction of more than three degrees.

(d) Clevis pins subject to load or moion and retained only by cotter pins may lot be used in the control system.

(e) Turnbuckles must be attached to arts having angular motion in a manner hat will positively prevent binding hroughout the range of travel.

(f) Tab control cables are not part of he primary control system and may be ss than 8 inch diameter in airplanes hat are safely controllable with the tabs 1 the most adverse positions.

§ 23.697 Wing flap controls.

(a) Each wing flap control must be designed so that, when the flap has been placed in any position upon which compliance with the performance requirements of this part is based, the flap will not move from that position unless the control is adjusted or is moved by the automatic operation of a flap load limiting device.

(b) The rate of movement of the flaps in response to the operation of the pilot's control or automatic device must give satisfactory flight and performance characteristics under steady or changing conditions of airspeed, engine power, and attitude.

§ 23.699 Wing flap position indicator. There must be a wing flap position indicator for

(a) Flap installations with only the retracted and fully extended position, unless

(1) A direct operating mechanism provides a sense of "feel" and position (such as when a mechanical linkage is employed); or

(2) The flap position is readily determined without seriously detracting from other piloting duties under any flight condition, day or night; and

(b) Flap installation with intermediate flap positions if

(1) Any flap position other than retracted or fully extended is used to show compliance with the performance requirements of this part; and

(2) The flap installation does not meet the requirements of paragraph (a) (1) of this section.

§ 23.701 Flap interconnection.

(a) The motion of flaps on opposite sides of the plane of symmetry must be synchronized by a mechanical interconnection unless the airplane has safe flight characteristics with the flaps retracted on one side and extended on the other.

(b) If an interconnection is used in multiengine airplanes, it must be designed to account for the unsymmetrical loads resulting from flight with the engines on one side of the plane of symmetry inoperative and the remaining engines at takeoff power. For singleengine airplanes, it may be assumed that 100 percent of the critical air load acts on one side and 70 percent on the other.

(a) The shock absorbing elements in main, nose, and tail wheel units must be substantiated by the tests specified in § 23.723.

(b) The shock absorbing ability of the landing gear during taxiing must be shown in the operational tests required by § 23.235.

§ 23.723 Shock absorption tests.

(a) It must be shown by energy absorption tests that the limit load factors selected for design under § 23.473 will not be exceeded in landings with the limit descent velocity specified in that section.

(b) The landing gear may not fail, but may yield, in a test showing its reserved energy absorption capacity, simulating a descent velocity of 1.2 times the limit descent velocity, assuming wing lift equal to the weight of the airplane.

§ 23.725 Limit drop tests.

(a) If compliance with § 23.723 (a) is shown by free drop tests, these tests must be made on the complete airplane, or on units consisting of wheel, tire, and shock absorber, in their proper relation, from free drop heights not less than those determined by the following formula:

h (inches) 3.6 (W/S) 1/2

However, the free drop height may not be less than 9.2 inches and need not be more than 18.7 inches.

(b) If wing lift is simulated in free drop tests, the landing gear must be dropped with an effective weight equal to ·h+(1—L) d

W=W

where

h+d

We the effective weight to be used in the drop test (lbs.);

h specified free drop height (inches); d=deflection under impact of the tire

(at the approved inflation pressure) plus the vertical component of the axle travel relative to the drop mass (inches);

W=WM for main gear units (lbs.), equal to the static weight on that unit with the airplane in the level attitude (with the nose wheel clear in the case of nose wheel type airplanes);

W=WT

for tail gear units (lbs.), equal to the static weight on the tail unit with the airplane in the tail-down attitude;

W=WN for nose wheel units (lbs.), equal to the vertical component of the static reaction that would exist af the nose wheel, assuming that the mass of the airplane acts at the center of gravity and exerts a force of 1.0g downward and 0.33g for ward; and

L=the ratio of the assumed wing lift t the airplane weight, but not mor than 0.667.

(c) The attitude in which a landing gear unit is drop tested must simulat the critical landing conditions for th unit.

(d) The value of d used in the com putation of We in paragraph (b) of thi section may not exceed the value actu ally obtained in the drop test.

(e) The limit inertia load factor mus be determined from the drop test i paragraph (b) of this section accordin to the following formula:

n=ny we + L

where

nj= the load factor developed in the dro test (that is, the acceleratio (dv/dt) in g's recorded in the dro test) plus 1.0; and

We, W, and L are the same as in the dro test computation.

(f) The value of n determined in ac cordance with paragraph (e) may no be more than the limit inertia load fac tor used in the landing conditions 1 § 23.473. § 23.727

Reserve energy absorptio drop test.

(a) If compliance with the reserve er ergy absorption requirement in § 23.72 (b) is shown by free drop tests, the dro height may not be less than 1.44 tim that specified in § 23.725.

(b) If wing lift equal to the airpla weight is simulated, the units must 1 dropped with an effective mass equal

We= =W other details are the same as in § 23.72

(hd), where the symbols a

§ 23.729 Retracting mechanism.

(a) General. For airplanes with r tractable landing gear, the followi apply:

(1) Each landing gear retracti mechanism and its supporting structu must be designed for maximum flig load factors with the gear retracted a must be designed for the combination friction, inertia, brake torque, and

loads, occurring during retraction at any airspeed up to 1.6 Vs, with flaps retracted, and for any load factor up to those specified in § 23.345 for the flapsextended condition.

(2) The landing gear and retracting mechanism, including the wheel well doors, must withstand flight loads with the landing gear extended at any speed up to at least 1.6 V with flaps retracted. $1 (b) Landing gear lock. There must be positive means (other than the use of hydraulic pressure) to keep the landing gear extended.

(c) Emergency operation. A landplane without manually operated landing gear must have an auxiliary means of extending the gear.

(d) Operation test. The proper functloning of the retracting mechanism must be shown by operation tests.

(e) Position indicator and warning device. There must be means to indicate to the pilot when the wheels are secured in the extreme positions. In addition, landplanes must have an aural or equally effective warning device that functions continuously, when one more throttles are closed, until the gear is down and locked. A throttle stop is not an acceptable alternative to an aural landing gear warning device.

§ 23.731 Wheels.

or

(a) Each main and nose wheel must be approved.

(b) The maximum static load rating of each wheel may not be less than the corresponding static ground reaction with

(1) Design maximum weight; and (2) Critical center of gravity.

(c) The maximum limit load rating of each wheel must equal or exceed the maximum radial limit load determined under the applicable ground load requirements of this part.

$23.733 Tires.

(a) Each landing gear wheel must have a tire

(1) That is a proper fit on the rim of the wheel; and

(2) Whose tire rating (assigned by the Tire and Rim Association or the Administrator) is not exceeded

(i) By a load on each main wheel tire equal to the corresponding static ground reaction under the design maximum weight and critical center of gravity; and

(ii) By a load on nose wheel tires (to be compared with the dynamic rating established for such tires) equal to the reaction obtained at the nose wheel, assuming the mass of the airplane to be concentrated at the most critical center of gravity and exerting a force of 1.0W downward and 0.31W forward (where W is the design maximum weight), with the reactions distributed to the nose and main wheels by the principles of statics, and with the drag reaction at the ground applied only at wheels with brakes.

(b) If specially constructed tires are used, the wheels must be plainly and conspicuously marked to that effect. The markings must include the make, size, number of plies, and identification marking of the proper tire.

§ 23.751

Main float buoyancy.

(a) Each main float must have

(1) A buoyancy of 80 percent in excess of that required to support the maximum weight of the seaplane or amphibian in fresh water; and

(2) Enough watertight compartments to provide reasonable assurance that the seaplane or amphibian will stay afloat if any two compartments of the main floats are flooded.

(b) Each main float must contain at least four watertight compartments approximately equal in volume.

§ 23.753 Main float design.

Each seaplane main float must be approved and must meet the requirements of § 23.521.

§ 23.755 Hulls.

(a) The hull of a hull seaplane or amphibian of 1,500 pounds or more maximum weight must have watertight com

partments designed and arranged so that the hull, auxiliary floats, and tires (if used), will keep the airplane afloat in fresh water when

(1) For airplanes of 5,000 pounds or more maximum weight, any two adjacent compartments are flooded; and

(2) For airplanes of 1,500 pounds up to, but not including, 5,000 pounds maximum weight, any single compartment is flooded.

(b) The hulls of hull seaplanes or amphibians of less than 1,500 pounds maximum weight need not be compartmented.

(c) Bulkheads with watertight doors may be used for communication between compartments.

§ 23.757 Auxiliary floats.

Auxiliary floats must be arranged so that, when completely submerged in fresh water, they provide a righting moment of at least 1.5 times the upsetting moment caused by the seaplane or amphibian being tilted.

PERSONNEL AND CARGO ACCOMMODATIONS § 23.771 Pilot compartment.

For each pilot compartment

(a) The compartment and its equipment must allow each pilot to perform his duties without unreasonable concentration or fatigue; and

(b) The aerodynamic controls listed in § 23.779, excluding cables and control rods, must be located with respect to the propellers so that no part of the pilot or the controls lies in the region between the plane of rotation of any inboard propeller and the suface generated by a line passing through the center of the propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. § 23.773

Pilot compartment view.

(a) Each pilot compartment must be free from glare and reflections that could interfere with the pilot's vision, and designed so that

(1) The pilot's view is sufficiently extensive, clear, and undistorted, for safe operation; and

(2) Each pilot is protected from the elements so that moderate rain conditions do not unduly impair his view of the flight path in normal flight and while landing.

(b) If certification for night operation is requested, compliance with paragraph (a) of this section must be shown in night flight tests.

§ 23.775 Windshields and windows.

(a) Nonsplintering safety glass must be used in internal glass panes.

(b) The design of windshields, win dows, and canopies in pressurized air planes must be based on factors peculia to high altitude operation, including(1) The effects of continuous and cyclic pressurization loadings;

(2) The inherent characteristics o the material used; and

(3) The effects of temperatures an temperature gradients.

(c) On pressurized airplanes, an in closure canopy including a representa tive part of the installation must b subjected to special tests to account fo the combined effects of continuous and cyclic pressurization loadings and fligh loads.

(d) The windshield and side window forward of the pilot's back when he 1 seated in the normal flight position mus have a luminous transmittance value o not less than 70 percent.

§ 23.777 Cockpit controls.

(a) Each cockpit control must be lo cated and (except where its function i obvious) identified to provide convenien operation and to prevent confusion and inadvertent operation.

(b) The controls must be located and arranged so that the pilot, when seated, has full and unrestricted movement of each control without interference from either his clothing or the cockpit struc ture.

(c) Identical powerplant controls fo each engine must be located to preven confusion as to the engines they control

(d) Wing flap and auxiliary lift devic controls must be located

(1) Centrally, or to the right of th pedestal or powerplant throttle contro centerline; and

(2) Far enough away from the land ing gear control to avoid confusion.

(e) The landing gear control must b located to the left of the throttle center line or pedestal centerline.

§ 23.779 Motion and effect of cockp controls.

Cockpit controls must be designed s that they operate as follows:

sponding to the specified flight and ground load conditions, including the emergency landing conditions prescribed in § 23.561.

(b) Each seat and berth must be approved.

(c) Each pilot seat must be designed for the reactions resulting from the application of pilot forces to the primary flight controls, as prescribed in § 23.395.

(d) Unless otherwise placarded, each seat in utility and acrobatic category airplanes must be designed to accommodate passengers wearing parachutes.

(e) Each berth installed parallel to the longitudinal axis of an airplane must be designed so that the forward part has a padded end-board, canvas diaphragm, or equivalent means that can withstand the static load reaction of the occupant when the occupant is subjected to the forward inertia forces prescribed in § 23.561. In addition

(1) The berth must have an approved safety belt and may not have corners or other parts likely to cause serious injury to a person occupying it during emergency conditions; and

(2) Safety belt attachments for the berth must be designed to withstand the critical loads resulting from relevant flight and ground load conditions and from the emergency landing conditions prescribed in § 23.561, with the exception of the forward load.

(f) Proof of compliance with the strength and deformation requirements of this section for seats and berths approved as part of the type design and for seat and berth installations may be shown by

(1) Structural analysis, if the structure conforms to conventional airplane types for which existing methods of analysis are known to be reliable;

(2) A combination of structural analysis and static load tests to limit loads;

or

(3) Static load tests to ultimate loads. The inertia forces prescribed in § 23.561 must be multiplied by a factor of 1.33 (rather than by the fitting factor prescribed in § 23.625) in determining the strength of the atachment of each seat or berth to the structure.

§ 23.787 Cargo compartments.

(a) Each cargo compartment must be designed for its placarded maximum weight of contents and for the critical load distributions at the appropriate