Part V  Airworthiness Manual Chapter 523  Normal, Utility, Aerobatic And Commuter Category Aeroplanes
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Canadian Aviation Regulations (CARs) 20172
Content last revised: 2010/06/01
SUBCHAPTERS
 A (523.1523.3),
 B (523.21523.253),
 C (523.301523.575),
 D (523.601523.871),
 E (523.901523.1203),
 F (523.1301523.1461),
 G (523.1501523.1589)
APPENDICES
(2002/03/01; no previous version)
SUBCHAPTER C STRUCTURE  GENERAL
523.301 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.

(b) Unless otherwise provided, the air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the aeroplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. Methods used to determine load intensities and distribution on canard and tandem wing configurations must be validated by flight test measurement unless the methods used for determining those loading conditions are shown to be reliable or conservative on the configuration under consideration.

(c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account.

(d) Simplified structural design criteria may be used if they result in design loads not less than those prescribed in 523.331 through 523.521. For aeroplane configurations described in appendix A, 523.1, the design criteria of appendix A of this chapter are an approved equivalent of 523.321 through 523.459. If appendix A of this chapter is used, the entire appendix must be substituted for the corresponding sections of this chapter.
(Change 5233 (920102))
(Change 5235)
523.302 Canard or Tandem Wing Configurations
The forward structure of a canard or tandem wing configuration must:

(a) Meet all requirements of subchapter C and subchapter D of this chapter applicable to a wing; and

(b) Meet all requirements applicable to the function performed by these surfaces.
(Change 5233 (920102))
523.303 Factor of Safety
Unless otherwise provided, a factor of safety of 1.5 must be used.
523.305 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.
(Change 5234 (960901))
523.307 Proof of Structure

(a) Compliance with the strength and deformation requirements of 523.305 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 Subchapter D of this Chapter.
Flight Loads
523.321 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 subchapter 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 limitations specified in 523.1583 through 523.1589.


(c) When significant, the effects of compressibility must be taken into account.
(Change 5234 (960901))
523.331 Symmetrical Flight Conditions

(a) The appropriate balancing horizontal tail load must be accounted for in a rational or conservative manner when determining the wing loads and linear inertia loads corresponding to any of the symmetrical flight conditions specified in 523.333 through 523.341.

(b) The incremental horizontal tail loads due to manoeuvring and gusts must be reacted by the angular inertia of the aeroplane in a rational or conservative manner.

(c) Mutual influence of the aerodynamic surfaces must be taken into account when determining flight loads.
(Change 5233 (920102))
523.333 Flight Envelope

(a) General. Compliance with the strength requirements of this subchapter 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 paragraph (d) of this section) that represents the envelope of the flight loading conditions specified by the manoeuvring and gust criteria of paragraphs (b) and (c) of this section 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 523.337 at speeds up to V_{D};

(2) The negative manoeuvring load factor specified in 523.337 at V_{C}; and

(3) Factors varying linearly with speed from the specified value at V_{C} to 0.0 at V_{D} for the normal and commuter category, and 1.0 at V_{D} 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 f.p.s. at V_{C} must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 50 f.p.s. at 20,000 feet to 25 f.p.s. at 50,000 feet.

(ii) Positive and negative gusts of 25 f.p.s. at V_{D} must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 25 f.p.s. at 20,000 feet to 12.5 f.p.s. at 50,000 feet.

(iii) In addition, for commuter category aeroplanes, positive (up) and negative (down) rough air gusts of 66 f.p.s. at V_{B} must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 66 f.p.s. at 20,000 feet to 38 f.p.s. at 50,000 feet.


(2) The following assumptions must be made:

(i) The shape of the gust is:
where:
s = Distance penetrated into gust (ft.);
C= Mean geometric chord of wing (ft.); and
U_{de} = Derived gust velocity referred to in subparagraph (l) of this section.

(ii) Gust Load factors vary linearly with speed between V_{C} and V_{D}.



(d) Flight envelope.
(Change 5231 (880101))
523.335 Design Airspeeds
Except as provided in paragraph(a)(4) of this section, the selected design airspeeds are equivalent airspeeds (EAS).

(a) Designcruising speed, V_{C}. For V_{C} the following apply:

(1) Where W/S' = wing loading at the design maximum takeoff weight, V_{C} (in knots) may not be less than:

(i) 33Ö(W/S) (for normal, utility, and commuter category aeroplanes);

(ii) 36Ö(W/S) (for aerobatic category aeroplanes).


(2) For values of W/S more than 20, the multiplying factors may be decreased linearly with W/S to a value of 28.6 where W/S = 100.

(3) V_{C} need not be more than 0.9 V_{H} at sea level.

(4) At altitudes where an M_{D} is established, a cruising speed M_{C} limited by compressibility may be selected.


(b) Design dive speed V_{D}. For V_{D}, the following apply:

(1) V_{D}/M_{D} may not be less than l.25 V_{C}/M_{C}; and

(2) With V_{C} min, the required minimum design cruising speed, V_{D} (in knots) may not be less than:

(i) 1.40 V_{C} min (for normal and commuter category aeroplanes);

(ii) l.50 V_{C} min (for utility category aeroplanes; and

(iii) l.55 V_{C} min (aerobatic category aeroplanes).


(3) For values of W/S more than 20, the multiplying factors in subparagraph (2) of this paragraph may be decreased linearly with W/S to a value of 1.35 where W/S = l00.

(4) Compliance with subparagraphs (1) and (2) of this paragraph need not be shown if V_{D}/M_{D} is selected so that the minimum speed margin between V_{C}/M_{C} and V_{D}/M_{D} is the greater of the following:

(i) The speed increase resulting when, from the initial condition of stabilised flight at V_{C}/M_{C}, the aeroplane is assumed to be upset, flown for 20 seconds along a flight path 7.5° below the initial path, and then pulled up with a load factor of 1.5 (0.5 g. acceleration increment). At least 75 percent maximum continuous power for reciprocating engines, and maximum cruising power for turbines, or, if less, the power required for V_{C}/M_{C} for both kinds of engines, must be assumed until the pull up is initiated, at which point power reduction and pilotcontrolled drag devices may be used; and either:

(ii) Mach 0.05 for normal, utility, and aerobatic category aeroplanes (at altitudes where M_{D} is established); or

(iii) Mach 0.07 for commuter category aeroplanes (at altitudes where M_{D} is established) unless a rational analysis, including the effects of automatic systems, is used to determine a lower margin. If a rational analysis is used, the minimum speed margin must be enough to provide for atmospheric variations (such as horizontal gusts), and the penetration of jet streams or cold fronts), instrument errors, airframe production variations, and must not be less than Mach 0.05.



(c) Design manoeuvring speed V_{A}. For V_{A} the following applies:

(1) V_{A} may not be less than V_{S} ,
where:

(i) V_{S} is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum aeroplane normal force coefficients, C_{NA}; and

(ii) n is the limit manoeuvring load factor used in design.


(2) The value of V_{A} need not exceed the value of V_{C} used in design.


(d) Designspeed for maximum gust intensity, V_{B}. For V_{B}, the following apply:

(1) V_{B} may not be less than the speed determined by the intersection of the line representing the maximum positive lift C_{N} _{MAX,} and the line representing the rough air gust velocity on the gust Vn diagram, or V_{S1} , whichever is less, where:

(i) n_{g} the positive aeroplane gust load factor due to gust, at speed V_{C} (in accordance with 523.341), and at the particular weight under consideration; and

(ii) V_{S1} is the stalling speed with the flaps retracted at the particular weight under consideration.


(2) V_{B} need not be greater than V_{C}.

(Change 5231 (880101))
(Change 5235)
523.337 Limit Manoeuvring Load Factors

(a) The positive limit manoeuvring load factor n may not be less than:

(1) 2.1+for normal and commuter category aeroplanes, where W =design maximum takeoff weight, except that n need not be more than 3.8;

(2) 4.4 for utility category aeroplanes; or

(3) 6.0 for aerobatic category aeroplanes.


(b) The negative limit manoeuvring load factor may not be less than:

(1) 0.4 times the positive load factor for the normal, utility, and commuter categories; or

(2) 0.5 times the positive load factor for the aerobatic category.


(c) Manoeuvring load factors lower than those specified in this section may be used if the aeroplane has design features that make it impossible to exceed these values in flight.
(Change 5231 (880101))
(Change 5235)
523.341 Gust Loads Factors

(a) Each aeroplane must be designed to withstand loads on each lifting surface resulting from gusts specified in 523.333(c).

(b) The gust load for a canard or tandem wing configuration must be computed using a rational analysis, or may be computed in accordance with paragraph (c) of this section, provided that the resulting net loads are shown to be conservative with respect to the gust criteria of 523.333(c).
(c) In the absence of a more rational analysis, the gust load factors must be computed as follows:
where:
K_{g} = = gust alleviation factor;
m_{g} = = aeroplane mass ratio;
U_{de} = Derived gust velocities referred to in 523.333(c) (f.p.s.);
? = Density of air (slugs/cu. ft.);
W/S = Wing loading (p.s.f.) due to the applicable weight of the aeroplane in the particular load case.
C= Mean geometric chord (ft.);
g = Acceleration due to gravity (ft/sec.^{2});
V = Aeroplane equivalent speed (knots); and
a = Slope of the aeroplane normal force coefficient curve C_{NA} per radian if the gust loads are applied to the wings and horizontal tail surfaces simultaneously by a rational method. The wing lift curve slope C_{L} per radian may be used when the gust load is applied to the wings only and the horizontal tail gust loads are treated as a separate condition.
(Change 5233 (920102))
(Change 5235)
523.343 Design Fuel Loads

(a) The disposable load combinations must include each fuel load in the range from zero fuel to the selected maximum fuel load.

(b) If fuel is carried in the wings, the maximum allowable weight of the aeroplane without any fuel in the wing tank(s) must be established as "maximum zero wing fuel weight," if it is less than the maximum weight.

(c) For commuter category aeroplanes, a structural reserve fuel condition, not exceeding fuel necessary for 45 minutes of operation at maximum continuous power, may be selected. If a structural reserve fuel condition is selected, it must be used as the minimum fuel weight condition for showing compliance with the flight load requirements prescribed in this chapter and:

(1) The structure must be designed to withstand a condition of zero fuel in the wing at limit loads corresponding to:

(i) Ninety percent of the manoeuvring load factors defined in 523.337, and

(ii) Gust velocities equal to 85 percent of the values prescribed in 523.333(c).


(2) The fatigue evaluation of the structure must account for any increase in operating stresses resulting from the design condition of paragraph(c)(1) of this section.

(3) The flutter, deformation, and vibration requirements must also be met with zero fuel in the wings.

(Change 5235)
523.345 High Lift Devices

(a) If flaps or similar high lift devices are to be used for takeoff, approach or landing, the aeroplane, with the flaps fully extended at V_{F}, is assumed to be subjected to symmetrical manoeuvres and gusts within the range determined by:

(1) Manoeuvring, to a positive limit load factor of 2.0; and

(2) Positive and negative gust of 25 feet per second acting normal to the flight path in level flight.


(b) V_{F} must be assumed to be not less than 1.4 V_{S} or 1.8 V_{SF}, whichever is greater, where:

(1) V_{S} is the computed stalling speed with flaps retracted at the design weight; and

(2) V_{SF} is the computed stalling speed with flaps fully extended at the design weight.

(3) If an automatic flap load limiting device is used, the aeroplane may be designed for the critical combinations of airspeed and flap position allowed by that device.


(c) In determining external loads on the aeroplane as a whole, thrust, slipstream, and pitching acceleration may be assumed to be zero.

(d) The flaps, their operating mechanism, and their supporting structures, must be designed to withstand the conditions prescribed in paragraph (a) of this section. In addition, with the flaps fully extended at V_{F}, the following conditions, taken separately, must be accounted for:

(1) A headon gust having a velocity of 25 feet per second (EAS), combined with propeller slipstream corresponding to 75 percent of maximum continuous power; and

(2) The effects of propeller slipstream corresponding to maximum takeoff power.

(Change 5235)
523.347 Unsymmetrical Flight Conditions

(a) The aeroplane is assumed to be subjected to the unsymmetrical flight conditions of 523.349 and 523.351. Unbalanced aerodynamic moments about the centre of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces.

(b) Aerobatic category aeroplanes certified for flick manoeuvres (snap roll) must be designed for additional asymmetric loads acting on the wing and the horizontal tail.
(Change 5235)
523.349 Rolling Conditions
The wing and wing bracing must be designed for the following loading conditions:

(a) Unsymmetrical wing loads appropriate to the category. Unless the following values result in unrealistic loads, the rolling accelerations may be obtained by modifying the symmetrical flight conditions in 523.333(d) as follows:

(1) For the aerobatic category, in conditions A and F, assume that 100 percent of the semispan wing air load acts on one side of the plane of symmetry and 60 percent of this load acts on the other side.

(2) For normal, utility, and commuter categories, in Condition A, assume that 100 percent of the semispan wing airload acts on one side of the aeroplane and 75 percent of this load acts on the other side.


(b) The loads resulting from the aileron deflections and speeds specified in 523.455, in combination with an aeroplane load factor of at least two thirds of the positive manoeuvring factor used for design. Unless the following values result in unrealistic loads, the effect of aileron displacement on wing torsion may be accounted for by adding the following increment to the basic airfoil moment coefficient over the aileron portion of the span in the aileron portion of the span in the critical condition determined in 523.333(d):
?C_{m} = 0.01d
where:
?C_{m} is the moment coefficient increment; and
d is the down aileron deflection in degrees in the critical condition.
(Change 5231 (880101))
(Change 5235)
523.351 Yawing Conditions
The aeroplane must be designed for yawing loads on the vertical surfaces resulting from the loads specified in 523.441 through 523.445.
(Change 5233 (920102))
523.361 Engine Torque

(a) Each engine mount and its supporting structure must be designed for the effects of:

(1) A limit engine torque corresponding to takeoff power and propeller speed acting simultaneously with 75 percent of the limit loads from flight condition A of 523.333(d);

(2) A limit engine torque corresponding to maximum continuous power and propeller speed acting simultaneously with the limit loads from flight condition A of 523.333(d); and

(3) For turbo propeller installations, in addition to the conditions specified in paragraphs (a)(l) and (a)(2) of this section, a limit engine torque corresponding to takeoff power and propeller speed, multiplied by a factor accounting for propeller control system malfunction, including quick feathering, acting simultaneously with 1g level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used.


(b) For turbine engine installations, the engine mounts and supporting structure must be designed to withstand each of the following:

(1) A limit engine torque load imposed by sudden engine stoppage due to malfunction or structural failure (such as compressor jamming).

(2) A limit engine torque load imposed by the maximum acceleration of the engine.


(c) The limit engine torque to be considered under paragraph (a) of this section must be obtained by multiplying the mean torque by a factor of:

(1) 1.25 for turbo propeller installations;

(2) 1.33 for engines with five or more cylinders; and

(3) Two, three, or four, for engines with four, three, or two cylinders respectively.

(Change 5234 (960901))
523.363 Side Load on Engine Mount

(a) Each engine mount and its supporting structure must be designed for a limit load factor in a lateral direction, for the side load on the engine mount, of not less than:

(1) 1.33; or

(2) Onethird of the limit load factor for flight condition A.


(b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions.
523.365 Pressurized Cabin Loads
For each pressurized compartment, the following apply:

(a) The aeroplane structure must be strong enough to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting.

(b) The external pressure distribution in flight, and any stress concentrations, must be accounted for.

(c) If landings may be made, with the cabin pressurised, landing loads must be combined with pressure differential loads from zero up to the maximum allowed during landing.

(d) The aeroplane structure must be strong enough to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33, omitting other loads.

(e) If a pressurized cabin has two or more compartments separated by bulkheads or a floor, the primary structure must be designed for the effects of sudden release of pressure in any compartment with external doors or windows. This condition must be investigated for the effects of failure of the largest opening in the compartment. The effects of intercompartmental venting may be considered.
523.367 Unsymmetrical Loads Due to Engine Failure

(a) Turbo propeller aeroplanes must be designed for the unsymmetrical loads resulting from the failure of the critical engine including the following conditions in combination with a single malfunction of the propeller drag limiting system, considering the probable pilot corrective action on the flight controls:

(1) At speeds between V_{MC} and V_{D}, the loads resulting from power failure because of fuel flow interruption are considered to be limit loads.

(2) At speeds between V_{MC} and V_{C}, 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 enginepropeller combination.

(4) The timing and magnitude of the probable pilot corrective action must be conservatively estimated, considering the characteristics of the particular enginepropelleraeroplane combination.


(b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than 2 seconds after the engine failure. The magnitude of the corrective action may be based on the limit pilot forces specified in 523.397 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.
523.369 Rear Lift Truss

(a) If a rear lift truss is used, it must be designed to withstand conditions of reversed airflow at a design speed of:
where:
W/S = wing loading at design maximum takeoff weight.

(b) Either aerodynamic data for the particular wing section used, or a value of C_{L} equalling 0.8 with a chord wise distribution that is triangular between a peak at the trailing edge and zero at the leading edge, must be used.
(Change 5234 (960901))
(Change 5235)
523.371 Gyroscopic and Aerodynamic Loads

(a) Each engine mount and its supporting structure must be designed for the gyroscopic, inertial, and aerodynamic loads that result, with the engine(s) and propeller(s), if applicable, at maximum continuous rpm, under either:

(b) For aeroplanes approved for aerobatic manoeuvres, each engine mount and its supporting structure must meet the requirements of paragraph (a) of this section and be designed to withstand the load factors expected during combined maximum yaw and pitch velocities.

(c) For aeroplanes certificated in the commuter category, each engine mount and its supporting structure must meet the requirements of paragraph (a) of this section and the gust conditions specified in 523.341 of this chapter.
(Change 5234 (960901))
(Change 5235)
523.373 Speed Control Devices
If speed control devices (such as spoilers and drag flaps) are incorporated for use in en route conditions:

(a) The aeroplane must be designed for the symmetrical manoeuvres and gusts prescribed in 523.333, 523.337, and 523.341, and the yawing manoeuvres and lateral gusts in 523.441 and 523.443, with the device extended at speeds up to the placard device extended speed; and

(b) If the device has automatic operating or load limiting features, the aeroplane must be designed for the manoeuvre and gust conditions prescribed in paragraph (a) of this section at the speeds and corresponding device positions that the mechanism allows.
Control Surface and System Loads
523.391 Control Surface Loads

(a) The control surface loads specified in 523.397 through 523.459 are assumed to occur in the conditions described in 523.331 through 523.351.

(b) Removed
(Change 5233 (920102))
(Change 5235)
523.393 Loads Parallel To Hinge Line

(a) Control surfaces and supporting hinge brackets must be designed to withstand inertial loads acting parallel to the hinge line.

(b) In the absence of more rational data, the inertial loads may be assumed to be equal to KW, where:

(1) K = 24 for vertical surfaces

(2) K = 12 for horizontal surfaces; and

(3) W = weight of the movable surfaces.

(Change 5235)
523.395 Control System Loads

(a) Each flight control system and its supporting structure must be designed for loads corresponding to at least 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in 523.391 through 523.459. In addition, the following apply:

(1) The system limit loads need not exceed the higher of the loads that can be produced by the pilot and automatic devices operating the controls. However, autopilot forces need not be added to pilot forces. The system must be designed for the maximum effort of the pilot or autopilot, whichever is higher. In addition, if the pilot and the autopilot act in opposition, the part of the system between them may be designed for the maximum effort of the one that imposes the lesser load. Pilot forces used for design need not exceed the maximum forces prescribed in 523.397(b).

(2) The design must, in any case, provide a rugged system for service use, considering jamming, ground gusts, taxiing downwind, control inertia, and friction. Compliance with this subparagraph may be shown by designing for loads resulting from application of the minimum forces prescribed in 523.397 (b).


(b) A 125 percent factor on computed hinge moments must be used to design elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be used if hinge moments are based on accurate flight test data, the exact reduction depending upon the accuracy and reliability of the data.

(c) Pilot forces used for design are assumed to act at the appropriate control grips or pads as they would in flight, and to react at the attachments of the control system to the control surface horns.
523.397 Limit Control Forces and Torques

(a) In the control surface flight loading condition, the air loads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (b) of this section. In applying this criterion, the effects of control system boost and servomechanisms, and the effects of tabs must be considered. The automatic pilot effort must be used for design if it alone can produce higher control surface loads than the human pilot.

(b) The limit pilot forces and torques are as in Table below.
Control Maximum Forces Or Torques For Design Weight, Weight Equal To Or Less Than 5,000Pounds^{1} Minimum Forces Or Torques^{2} Aileron:
Stick
Wheel^{3}
Elevator:
Stick
Wheel (symmetrical)
Wheel (unsymmetrical)^{5}
Rudder
67 lbs
50 d in.lbs.^{4}
167 lbs
200 lbs
.........................
200 lbs
40 lbs.
40 d in.lbs.^{4}
100 lbs.
100 lbs.
100 lbs.
150 lbs.^{1}For design weight (W) more than 5,000 pounds, the specified maximum values must be increased linearly with weight to 1.18 times the specified values at a design mass weight of 5700 kg (12566 lbs.) and for commuter category aeroplanes, the specified values must be increased linearly with weight to 1.35 times the specified values at a design weight of 19,000 pounds.
^{2}If the design of any individual set of control systems or surfaces makes these specified minimum forces or torques inapplicable, values corresponding to the present hinge moments obtained under sec. 523.415, but not less than 0.6 of the specified minimum forces or torques, may be used.
^{3}The critical parts of the aileron control system must also be designed for a single tangential force with a limit value of 1.25 times the couple force determined from the above criteria.
^{4}D = wheel diameter (inches).
^{5}The unsymmetrical force must be applied at one of the normal handgrip points on the control wheel.
(Change 5231 (880101))
(Change 5234 (960901))
523.399 Dual Control System

(a) Each dual control system must be designed to withstand the force of the pilots operating in opposition, using individual pilot forces not less than the greater of:

(1) 0.75 times those obtained under 523.395; or

(2) The minimum forces specified in 523.397(b).


(b) Each dual control system must be designed to withstand the force of the pilots applied together, in the same direction, using individual pilot forces not less than 0.75 times those obtained under 523.395.
(Change 5235)
523.405 Secondary Control System
Secondary controls, such as wheel brakes, spoilers, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls.
523.407 Trim Tab Effects
The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort. In these cases, the tabs are considered to be deflected in the direction that would assist the pilot. These deflections must correspond to the maximum degree of "out of trim" expected at the speed for the condition under consideration.
523.409 Tabs
Control surface tabs must be designed for the most severe combination of airspeed and tab deflection likely to be obtained within the flight envelope for any usable loading condition.
523.415 Ground Gust Conditions

(a) The control system must be investigated as follows for control surface loads due to ground gusts and taxiing downwind:

(1) If an investigation of the control system for ground gust loads is not required by subparagraph (2) of this paragraph, but the applicant elects to design a part of the control system for these loads, these loads need only be carried from control surface horns through the nearest stops or gust locks and their supporting structures.

(2) If pilot forces less than the minimums specified in 523.397(b) are used for design, the effects of surface loads due to ground gusts and taxiing downwind must be investigated for the entire control system according to the formula:
where:
H = limit hinge moment (ft.lbs.);
c = mean chord of the control surface aft of the hinge line (ft.);
S = area of control surface aft of the hinge line (sq. ft.);
q = dynamic pressure (p.s.f.) based on a design speed not less than 14.6 + 14.6 (f.p.s.) where W/S = wing loading at design maximum weight, except that the design speed need not exceed 88 (f.p.s.);
K = limit hinge moment factor for ground gusts derived in paragraph (b) of this section. (For ailerons and elevators, a positive value of K indicates a moment tending to depress the surface and a negative value of K indicates a moment tending to raise the surface).


(b) The limit hinge moment factor K for ground gusts must be derived as in Table below.
Surface K Position Of Controls (a) Aileron
(b) Aileron
(c) Elevator
(d) Elevator
(e) Rudder
(f) Rudder0.75
+0.50
+0.75
+0.75
+0.75
+0.75(a) Control column locked or lashed in midposition.
(b) Ailerons at full throw; + moment on one aileron, moment on the other.
(c) Elevator full up ().
(d) Elevator full down (+)
(e) Rudder in neutral
(f) Rudder at full throw. 
(c) At all weights between the empty weight and the maximum weight declared for tiedown stated in the appropriate manual, any declared tiedown points and surrounding structure, control system, surfaces and associated gust locks, must be designed to withstand the limit load conditions that exist when the aeroplane is tied down and that result from wind speeds of up to 65 knots horizontally from any direction.
(Change 5231 (880101))
(Change 5234 (960901))
(Change 5235)
Horizontal Stabilizing and Balancing Surface
523.421 Balancing Loads

(a) A horizontal surface balancing load is a load necessary to maintain equilibrium in any specified flight condition with no pitching acceleration.

(b) Horizontal balancing surfaces must be designed for the balancing loads occurring at any point on the limit manoeuvring envelope and in the flap conditions specified in 523.345.
(Change 5233 (920102))
523.423 Manoeuvring Loads
Each horizontal surface and its supporting structure, and the main wing of a canard or tandem wing configuration, must be designed for manoeuvring loads imposed by the following conditions:

(a) A sudden movement of the pitching control, at V_{A}, to the maximum aft movement, and the maximum forward movement, as limited by the control stops, or pilot effort, whichever is critical.

(b) A sudden aft movement of the pitching control, at speeds above V_{A}, followed by a forward movement of the pitching control resulting in the following combinations of the normal and angular acceleration (see Table below).
Condition Normal Acceleration (n) Angular Acceleration (radian / sec.^{2}) Noseup pitching 1.0 Nosedown pitching N_{m} Where:

(1) n_{m} = positive limit manoeuvring load factor used in the design of the aeroplane; and

(2) V = initial speed in knots.
The conditions in this paragraph involve loads corresponding to the loads that may occur in a "checked manoeuvre" (a manoeuvre in which the pitching control is suddenly displaced in one direction and then suddenly moved in the opposite direction). The deflections and timing of the "checked manoeuvre" must avoid exceeding the limit manoeuvring load factor. The total horizontal surface load for both noseup and nosedown pitching conditions is the sum of the balancing loads at V and the specified value of the normal load factor n, plus the manoeuvring load increment due to the specified value of the angular acceleration.

(Change 5231 (920102))
523.425 Gust Loads

(a) Each horizontal surface other than a main wing must be designed for loads resulting from:

(1) Gust velocities specified in 523.333 (c) with flaps retracted; and

(2) Positive and negative gusts of 25 f.p.s. nominal intensity at V_{F} corresponding to the flight conditions specified in 523.345 (a)(2).


(b) (Reserved)

(c) When determining the total load on the horizontal surfaces for the conditions specified in paragraph (a) of this section, the initial balancing loads for steady unaccelerated flight at the pertinent design speeds V_{F}, V_{C} and V_{D} must first be determined. The incremental load resulting from the gusts must be added to the initial balancing load to obtain the total load.

(d) In the absence of a more rational analysis, the incremental load due to the gust must be computed as follows only on aeroplane configurations with aftmounted, horizontal surfaces, unless its use elsewhere is shown to be conservative:
where:
?L_{ht} = Incremental horizontal tail load (lbs.);
K_{g} = Gust alleviation factor defined in section 523.341;
U_{de} = Derived gust velocity (f.p.s.)
V = Aeroplane equivalent speed (knots);
a_{ht} =Slope of aft horizontal tail lift curve (per radian);
S_{ht} =Area of horizontal lift surface (ft.^{2}); and
= Downwash factor.
(Change 5233 (920102))
523.427 Unsymmetrical Loads

(a) Horizontal surfaces other than main wing and their supporting structure must be designed for unsymmetrical loads arising from yawing and slipstream effects, in combination with the loads prescribed for the flight conditions set forth in 523.421 through 523.425.

(b) In the absence of more rational data for aeroplanes that are conventional in regard to location of engines, wings, horizontal surfaces other than main wing, and fuselage shape:

(1) 100 percent of the maximum loading from the symmetrical flight conditions may be assumed on the surface on one side of the plane of symmetry; and

(2) The following percentage of that loading must be applied to the opposite side % = 10010 (n1), where n is the specified positive manoeuvring load factor, but this value may not be more than 80 percent.


(c) For aeroplanes that are not conventional (such as aeroplanes with horizontal surfaces other than main wing having appreciable dihedral or supported by the vertical tail surfaces) the surfaces and supporting structures must be designed for combined vertical and horizontal surface loads resulting from each prescribed flight condition taken separately.
(Change 5233 (920102))
Vertical Surfaces
523.441 Manoeuvring Loads

(a) At speeds up to V_{A}, the vertical surfaces must be designed to withstand the following conditions. In computing the loads, the yawing velocity may be assumed to be zero:

(1) With the aeroplane in unaccelerated flight at zero yaw, it is assumed that the rudder control is suddenly displaced to the maximum deflection, as limited by the control stops or by limit pilot forces.

(2) With the rudder deflected as specified in paragraph (a)(1) of this section, it is assumed that the aeroplane yaws to the overswing sideslip angle. In lieu of a rational analysis, an overswing angle equal to 1.5 times the static sideslip angle of paragraph (a)(3) of this section may be assumed.

(3) A yaw angle of 15 degrees with the rudder control maintained in the neutral position (except as limited by pilot strength).


(b) For commuter category aeroplanes, the loads imposed by the following additional manoeuvre must be substantiated at speeds from V_{A} to V_{D}/M_{D}. When computing the tail loads:

(1) The aeroplane must be yawed to the largest attainable steady state sideslip angle, with the rudder at maximum deflection caused by any one of the following:

(i) Control surface stops;

(ii) Maximum available booster effort;

(iii) Maximum pilot rudder force as shown below:


(2) The rudder must be suddenly displaced from the maximum deflection to the neutral position.


(c) The yaw angles specified in paragraph (a)(3) of this section may be reduced if the yaw angle chosen for a particular speed cannot be exceeded in:

(1) Steady slip conditions;

(2) Uncoordinated rolls from steep banks; or

(3) Sudden failure of the critical engine with delayed corrective action.

(Change 5233 (920102))
(Change 5235)
523.443 Gust Loads

(a) Vertical surfaces must be designed to withstand, in unaccelerated flight at speed V_{C}, lateral gusts of the values prescribed for V_{C} in 523.333(c).

(b) In addition, for commuter category aeroplanes, the aeroplane is assumed to encounter derived gusts normal to the plane of symmetry while in unaccelerated flight at V_{B}, V_{C}, V_{D}, and V_{F}. The derived gusts and aeroplane speeds corresponding to these conditions, as determined by 523.341 and 523.345, must be investigated. The shape of the gust must be as specified in 523.333(c)(2)(i).

(c) In the absence of a more rational analysis, the gust load must be computed as follows:
Where:
L_{vt}= Vertical surface loads (lbs.);
gust alleviation factor;
= lateral mass ratio
U_{de} = Derived gust velocity (f.p.s.);
r = Air density (slugs/cu. ft.);
W = The applicable weight of the aeroplane in the particular load case (lbs.);
S_{vt} = Area of vertical surface (ft.^{2});
= Mean geometric chord of vertical surface (ft.);
a_{vt} = Lift curve slope of vertical surface (per radian);
K = Radius of gyration in yaw (ft.);
1_{vt} = Distance from aeroplane c.g. to lift centre of vertical surface (ft.);
g = Acceleration due to gravity (ft./sec.^{2}); and
V = Equivalent airspeed (knots).
(Change 5231 (880101))
(Change 5233 (920102))
(Change 5235)
523.445 Outboard Fins or Winglets

(a) If outboard fins or winglets are included on the horizontal surfaces or wings, the horizontal surfaces or wings must be designed for their maximum load in combination with loads induced by the fins or winglets and moments or forces exerted on the horizontal surfaces or wings by the fins or winglets.

(b) If outboard fins or winglets extend above and below the horizontal surface, the critical vertical surface loading (the load per unit area as determined under 523.441 and 523.443) must be applied to:

(1) The part of the vertical surfaces above the horizontal surface with 80 percent of that loading applied to the part below the horizontal surface; and

(2) The part of the vertical surfaces below the horizontal surface with 80 percent of that loading applied to the part above the horizontal surface.


(c) The end plate effects of outboard fins or winglets must be taken into account in applying the yawing conditions of 523.441 and 523.443 to the vertical surfaces in paragraph (b) of this section.

(d) When rational methods are used for computing loads, the manoeuvring loads of 523.441 on the vertical surfaces and the oneg horizontal surface load, including induced loads on the horizontal surface and moments or forces exerted on the horizontal surfaces by the vertical surfaces, must be applied simultaneously for the structural loading condition.
(Change 5231 (880101))
(Change 5233 (920102))
Ailerons and Special Devices
523.455 Ailerons

(a) The ailerons must be designed for the loads to which they are subjected:

(1) In the neutral position during symmetrical flight conditions; and

(2) By the following deflections (except as limited by pilot effort), during unsymmetrical flight conditions; and

(i) Sudden maximum displacement of the aileron control at V_{A}. Suitable allowance may be made for control system deflections.

(ii) Sufficient deflection at V_{C}, where V_{C} is more than V_{A}, to produce a rate of roll not less than obtained in subparagraph (2)(i).

(iii) Sufficient deflection at V_{A} to produce a rate of roll not less than onethird of that obtained in subparagraph (2)(i).



(b) (Reserved)
(Change 5233 (920102))
(Change 5235)
523.457 Wing Flaps [Removed]
(Change 5235)
523.459 Special Devices
The loading for special devices using aerodynamic surfaces (such as slots and spoilers) must be determined from test data.
Ground Loads
523.471 General
The limit ground loads specified in this subchapter 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.
523.473 Ground Load Conditions and Assumptions

(a) The ground load requirements of this subchapter must be complied with at the design maximum weight except that 523.479, 523.481, and 523.483 may be complied with at a design landing weight (the highest weight for landing conditions at the maximum descent velocity) allowed under paragraphs (b) and (c) of this section.

(b) The design landing weight may be as low as:

(1) 95 percent of the maximum weight if the minimum fuel capacity is enough for at least onehalf hour of operation at maximum continuous power plus a capacity equal to a fuel weight which is the difference between the design maximum weight and the design landing weight; or

(2) The design maximum weight less the weight of 25 percent of the total fuel capacity.


(c) The design landing weight of a multiengine aeroplane may be less than that allowed under paragraph (b) of this section if:

(1) The aeroplane meets the oneengineinoperative climb requirements of 523.67(b)(1) or (c); and

(2) Compliance is shown with the fuel jettisoning system requirements of 523.1001.


(d) The selected limit vertical inertia load factor at the centre of gravity of the aeroplane for the ground load conditions prescribed in this subchapter may not be less than that which would be obtained when landing with a descent velocity (V), in feet per second, equal to 4.4 (W/S)1/4, except that this velocity need not be more than 10 feet per second and may not be less than seven feet per second.

(e) Wing lift not exceeding twothirds of the weight of the aeroplane may be assumed to exist throughout the landing impact and to act through the centre of gravity. The ground reaction load factor may be equal to the inertia load factor minus the ratio of the above assumed wing lift to the aeroplane weight.

(f) If energy absorption tests are made to determine the limit load factor corresponding to the required limit descent velocities, these tests must be made under 523.723(a)

(g) No inertia load factor used for design purposes may be less than 2.67, nor may the limit ground reaction load factor be less than 2.0 at design maximum weight, unless these lower values will not be exceeded in taxiing at speeds up to takeoff speed over terrain as rough as that expected in service.
(Change 5234 (960901))
(Change 5235)
523.477 Landing Gear Arrangement
Sections 523.479 through 523.483, or the conditions in Appendix C, apply to aeroplanes with conventional arrangements of main and nose gear, or main and tail gear.
523.479 Level Landing Conditions

(a) For a level landing, the aeroplane is assumed to be in the following attitudes:

(1) For aeroplanes with tail wheels, a normal level flight attitude.

(2) For aeroplanes with nose wheels, attitudes in which:

(i) The nose and main wheels contact the ground simultaneously; and

(ii) The main wheels contact the ground and the nose wheel is just clear of the ground.
The attitude used in subdivision (i) of this subparagraph may be used in the analysis required under subdivision (ii) of this subparagraph.



(b) When investigating landing conditions, the drag components simulating the forces required to accelerate the tires and wheels up to the landing speed (spinup) must be properly combined with the corresponding instantaneous vertical ground reactions, and the forwardacting horizontal loads resulting from rapid reduction of the spinup drag loads (springback) must be combined with vertical ground reactions at the instant of the peak forward load, assuming wing lift and a tiresliding coefficient of friction of 0.8. However, the drag loads may not be less than 25 percent of the maximum vertical ground reactions (neglecting wing lift).

(c) In the absence of specific tests or a more rational analysis for determining the wheel spinup and springback loads for landing conditions, the method set forth in Appendix D of this chapter must be used. If Appendix D of this chapter is used, the drag components used for design must not be less than those given by Appendix C of this chapter.

(d) For aeroplanes with tip tanks or large overhung masses (such as turbopropeller or jet engines) supported by the wing, the tip tanks and the structure supporting the tanks or overhung masses must be designed for the effects of dynamic responses under the level landing conditions of either paragraph (a)(1) or (a)(2)(ii) of this section. In evaluating the effects of dynamic response, an aeroplane lift equal to the weight of the aeroplane may be assumed.
(Change 5234 (960901))
523.481 Tail Down Landing Conditions

(a) For a tail down landing, the aeroplane is assumed to be in the following attitudes:

(1) For aeroplanes with tail wheels, an attitude in which the main and tail wheels contact the ground simultaneously.

(2) For aeroplanes with nose wheels, a stalling attitude, or the maximum angle allowing ground clearance by each part of the aeroplane, whichever is less.


(b) For aeroplanes with either tail or nose wheels, ground reactions are assumed to be vertical, with the wheels up to speed before the maximum vertical load is attained.
523.483 Onewheel Landing Conditions
For the onewheel landing condition, the aeroplane is assumed to be in the level attitude and to contact the ground on one side of the main landing gear. In this attitude, the ground reactions must be the same as those obtained on that side under 523.479.
523.485 Side Load Conditions

(a) For the side load condition, the aeroplane is assumed to be in a level attitude with only the main wheels contacting the ground and with the shock absorbers and tires in their static positions.

(b) The limit vertical load factor must be 1.33, with the vertical ground reaction divided equally between the main wheels.

(c) The limit side inertia factor must be 0.83, with the side ground reaction divided between the main wheels so that:

(1) 0.5 (W) is acting inboard on one side; and

(2) 0.33 (W) is acting outboard on the other side.


(d) The side loads prescribed in paragraph (c) of this section are assumed to be applied at the ground contact point and the drag loads may be assumed to be zero.
(Change 5234 (960901))
523.493 Braked Roll Conditions
Under braked roll conditions, with the shock absorbers and tires in their static positions, the following apply:

(a) The limit vertical load factor must be 1.33.

(b) The attitudes and ground contacts must be those described in 523.479 for level landings.

(c) A drag reaction equal to the vertical reaction at the wheel multiplied by a coefficient of friction of 0.8 must be applied at the ground contact point of each wheel with brakes, except that the drag reaction need not exceed the maximum value based on limiting brake torque.
523.497 Supplementary Conditions for Tail Wheels
In determining the ground loads on the tail wheel and affected supporting structures, the following apply:

(a) For the obstruction load, the limit ground reaction obtained in the tail down landing condition is assumed to act up aft through the axle at 45 degrees. The shock absorber and tire may be assumed to be in their static positions.

(b) For the side load, a limit vertical ground reaction equal to the static load on the tail wheel, in combination with a side component of equal magnitude, is assumed. In addition:

(1) If a swivel is used, the tail wheel is assumed to be swivelled 90° to the aeroplane longitudinal axis with the resultant ground load passing through the axle;

(2) If a lock steering device, or shimmy damper is used, the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point; and

(3) The shock absorber and tire are assumed to be in their static positions.


(c) If a tail wheel, bumper, or an energy absorption device is provided to show compliance with 523.925(b), the following apply:

(1) Suitable design loads must be established for the tail wheel, bumper, or energy absorption device; and

(2) The supporting structure of the tail wheel, bumper, or energy absorption device must be designed to withstand the loads established in paragraph (c)(1) of this section.

(Change 5235)
523.499 Supplementary Conditions for Nose Wheels
In determining the ground loads on nose wheels and affected supporting structures, and assuming that the shock absorbers and tires are in their static positions, the following conditions must be met:

(a) For aft loads, the limit force components at the axle must be:

(1) A vertical component of 2.25 times the static load on the wheel; and

(2) A drag component of 0.8 times the vertical load.


(b) For forward loads, the limit force components at the axle must be:

(1) A vertical component of 2.25 times the static load on the wheel; and

(2) A forward component of 0.4 times the vertical load.


(c) For side loads, the limit force components at ground contact must be:

(1) A vertical component of 2.25 times the static load on the wheel; and

(2) A side component of 0.7 times the vertical load.


(d) For aeroplanes with a steerable nose wheel that is controlled by hydraulic or other power, at design takeoff weight with the nose wheel in any steerable position, the application of 1.33 times the full steering torque combined with a vertical reaction equal to 1.33 times the maximum static reaction on the nose gear must be assumed. However, if a torque limiting device is installed, the steering torque can be reduced to the maximum value allowed by that device.

(e) For aeroplanes with a steerable nose wheel that has a direct mechanical connection to the rudder pedals, the mechanism must be designed to withstand the steering torque for the maximum pilot forces specified in 523.397(b).
(Change 5235)
523.505 Supplementary Conditions for Skiplanes
In determining ground loads for skiplanes and assuming that the aeroplane is resting on the ground with one main ski frozen at rest and the other skis free to slide, a limit side force equal to 0.036 times the design maximum weight must be applied near the tail assembly, with a factor of safety of 1.
523.507 Jacking Loads

(a) The aeroplane must be designed for the loads developed when the aircraft is supported on jacks at the design maximum weight assuming he following load factors for landing gear jacking points at a three point attitude and for primary flight structure jacking points in the level attitude:

(1) Verticalload factor of 1.35 times the static reactions.

(2) Fore, aft, and lateral load factors of 0.4 times the vertical static reactions.


(b) The horizontal loads at the jack points must be reacted by inertia forces so as to result in no change in the direction of the resultant loads at the jack points.

(c) The horizontal loads must be considered in all combinations with the vertical load.
523.509 Towing Loads
The towing loads of this section must be applied to the design of tow fittings and their immediate attaching structure.

(a) The towing loads specified in paragraph (d) of this section must be considered separately. These loads must be applied at the towing fittings and must act parallel to the ground. In addition:

(1) A vertical load factor equal to 1.0 must be considered acting at the centre of gravity; and

(2) The shock struts and tires must be in their static positions.


(b) For towing points not on the landing gear but near the plane of symmetry of the aeroplane, the drag and side tow load components specified for the auxiliary gear apply. For towing points located outboard of the main gear, the drag and side tow load components specified for the main gear apply. Where the specified angle of swivel cannot be reached, the maximum obtainable angle must be used.

(c) The towing loads specified in paragraph (d) of this section must be reacted as follows:

(1) The side component of the towing load at the main gear must be reacted by a side force at the static ground line of the wheel to which the load is applied.

(2) The towing loads at the auxiliary gear and the drag components of the towing loads at the main gear must be reacted as follows:

(i) A reaction with a maximum value equal to the vertical reaction must be applied at the axle of the wheel to which the load is applied. Enough aeroplane inertia to achieve equilibrium must be applied.

(ii) The loads must be reacted by aeroplane inertia.



(d) The prescribed towing loads are as per Table below, where W is the design maximum weight:
Tow Point  Position  Load  

Magnitude  No.  Direction  
Main gear  0.225 W per main gear unit 
1 2 3 4 
Forward, parallel to drag axis. Forward, at 30° to drag axis. Aft, parallel to drag axis. Aft, at 30° to drag axis. 

Auxiliary gear 
Swivelled forward Swivelled aft 
0.3 W 
5 6 7 8 
Forward Aft Forward Aft 
Swivelled 45°from forward Swivelled 45° from aft 
0.15W 0.15W 
9 10 11 12 
Forward, in plane of wheel Aft, in plane of wheel Forward, in plane of wheel Aft, in plane of wheel 
523.511 Ground Load; Unsymmetrical Loads on MultipleWheel Units

(a) Pivoting loads. The aeroplane is assumed to pivot about on side of the main gear with:

(1) The brakes on the pivoting unit locked; and

(2) Loads corresponding to a limit vertical load factor of 1, and coefficient of friction of 0.8 applied to the main gear and its supporting structure.


(b) Unequal tire loads. The loads established under 523.471 through 523.483 must be applied in turn, in a 60/40 percent distribution, to the dual wheels and tires in each dual wheel landing gear unit.

(c) Deflated tire loads. For the deflated tire condition:

(1) 60 percent of the loads established under 523.471 through 523.483 must be applied in turn to each wheel in a landing gear unit; and

(2) 60 percent of the limit drag and side loads, and 100 percent of the limit vertical load established under 523.485 and 523.493 or lesser vertical load obtained under subparagraph (1) of this paragraph, must be applied in turn to each wheel in the dual wheel landing gear unit.

Water Loads
523.521 Water Load Conditions

(a) The structure of seaplanes and amphibians must be designed for water loads developed during takeoff and landing with the seaplane in any attitude likely to occur in normal operation at appropriate forward and sinking velocities under the most severe sea conditions likely to be encountered.

(b) Unless the applicant makes a rational analysis of the water loads, 523.523 through 523.537 apply.

(c) Removed
(Change 5234 (960901))
(Change 5235)
523.523 Design Weights and Centre of Gravity Positions

(a) Design weights. The water load requirements must be met at each operating weight up to the design landing weight except that, for the takeoff condition prescribed in 523.531, the design water takeoff weight (the maximum weight for water taxi and takeoff run) must be used.

(b) Centre of gravity positions. The critical centres of gravity within the limits for which certification is requested must be considered to reach maximum design loads for each part of the seaplane structure.
(Change 5234 (960901))
523.525 Application of Loads

(a) Unless otherwise prescribed, the seaplane as a whole is assumed to be subjected to the loads corresponding to the load factors specified in 523.527.

(b) In applying the loads resulting from the load factors prescribed in 523.527, the loads may be distributed over the hull or main float bottom (in order to avoid excessive local shear loads and bending moments at the location of water load application) using pressures not less than those prescribed in 523.533(b).

(c) For twin float seaplanes, each float must be treated as an equivalent hull on a fictitious seaplane with a weight equal to onehalf the weight of the twin float seaplane.

(d) Except in the takeoff condition of 523.531, the aerodynamic lift on the seaplane during the impact is assumed to be 2/3 of the weight of the seaplane.
(Change 5234 (960901))
523.527 Hull and Main Float Load Factors

(a) Water reaction load factors n_{w} must be computed in the following manner:

(1) For the step landing case

(2) For the bow and stern landing cases


(b) The following values are used:

(1) n_{w} = water reaction load factor (that is, the water reaction divided by seaplane weight).

(2) C_{1} = empirical seaplane operations factor equal to 0.012 (except that this factor may not be less than that necessary to obtain the minimum value of step load factor of 2.33).

(3) V_{SO} = seaplane stalling speed in knots with flaps extended in the appropriate landing position and with no slipstream effect.

(4) b = Angle of dead rise at the longitudinal station at which the load factor is being determined in accordance with figure Ii of Appendix I of this chapter.

(5) W = seaplane design landing weight in pounds.

(6) K_{1} = empirical hull station weighing factor, in accordance with figure Iii of Appendix I of this chapter.

(7) r_{x} = ratio of distance, measured parallel to hull reference axis, from the centre of gravity of the seaplane to the hull longitudinal station at which the load factor is being computed to the radius of gyration in pitch of the seaplane, the hull reference axis being a straight line, in the plane of symmetry, tangential to the keel at the main step.


(c) For a twin float seaplane, because of the effect of flexibility of the attachment of the floats to the seaplane, the factor K_{1} may be reduced at the bow and stern to 0.8 of the value shown in figure Iii of Appendix I of this chapter. This reduction applies only to the design of the carrythrough and seaplane structure.
(Change 5234 (960901))
523.529 Hull and Main Float Landing Conditions

(a) Symmetrical step, bow, and stern landing. For symmetrical step, bow, and stern landings, the limit water reaction load factors are those computed under 523.527. In addition:

(1) For symmetrical step landings, the resultant water load must be applied at the keel, through the centre of gravity, and must be directed perpendicularly to the keel line;

(2) For symmetrical bow landings, the resultant water load must be applied at the keel, onefifth of the longitudinal distance from the bow to the step, and must be directed perpendicularly to the keel line; and

(3) For symmetrical stern landings, the resultant water load must be applied at the keel, at a point 85 percent of the longitudinal distance from the step to the stern post, and must be directed perpendicularly to the keel line.


(b) Unsymmetrical landing for hull and single float seaplanes. Unsymmetrical step, bow, and stern landing conditions must be investigated. In addition:

(1) The loading for each condition consists of an upward component and a side component equal, respectively, to 0.75 and 0.25 tan b times the resultant load in the corresponding symmetrical landing condition; and

(2) The point of application and direction of the upward component of the load is the same as that in the symmetrical condition, and the point of application of the side component is at the same longitudinal station as the upward component but is directed inward perpendicularly to the plane of symmetry at a point midway between the keel and chine lines.


(c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading consists of an upward load at the step of each float of 0.75 and a side load of 0.25 tan b at one float times the step landing load reached under 523.527. The side load is directed inboard, perpendicularly to the plane of symmetry midway between the keel and chine lines of the float, at the same longitudinal station as the upward load.
(Change 5234 (960901))
523.531 Hull and Main Float Takeoff Condition
For the wing and its attachment to the hull or main float:

(a) The aerodynamic wing lift is assumed to be zero; and

(b) A downward inertia load, corresponding to a load factor computed from the following formula, must be applied:
where:
n = inertia load factor;
C_{TO} = empirical seaplane operations factor equal to 0.004;
V_{S1} = seaplane stalling speed (knots) at the design takeoff weight with the flaps extended in the appropriate takeoff position;
ß = angle of dead rise at the main step (degrees); and
W = design water takeoff weight in pounds.
(Change 5234 (960901))
523.533 Hull and Main Float Bottom Pressures

(a) General. The hull and main float structure, including frames and bulkheads, stringers, and bottom plating, must be designed under this section.

(b) Local pressures. For the design of the bottom plating and stringers and their attachments to the supporting structure, the following pressure distributions must be applied:

(1) For an unflared bottom, the pressure at the chine is 0.75 times the pressure at the keel, and the pressures between the keel and chine vary linearly, in accordance with figure Iiii of Appendix I of this chapter. The pressure at the keel (p.s.i.) is computed as follows:
where:
P_{k} = pressure (p.s.i.) at the keel;
C_{2} = 0.00213;
K_{2} = hull station weighing factor, in accordance with figure Iii of Appendix I of this chapter;
V_{S1} = seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and
b_{K} = angle of dead rise at keel, in accordance with figure Ii of Appendix I of this chapter.

(2) For a flared bottom, the pressure at the beginning of the flare is the same as that for an unflared bottom, and the pressure between the chine and the beginning of the flare varies linearly, in accordance with figure Iiii of Appendix I of this chapter. The pressure distribution is the same as that prescribed in paragraph (b)(1) of this section for an unflared bottom except that the pressure at the chine is computed as follows:
where:
P_{ch} = pressure (p.s.i.) at the chine;
C_{3} = 0.00l6;
K_{2} = hull station weighing factor, in accordance with figure Iii of Appendix I of this chapter;
V_{S1} = seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and
b = angle of dead rise at appropriate station.
The area over which these pressures are applied must simulate pressures occurring during high localized impacts on the hull or float, but need not extend over an area that would induce critical stresses in the frames or in the overall structure.


(c) Distributed pressures. For the design of the frames, keel, and chine structure, the following pressure distributions apply:

(1) Symmetrical pressures are computed as follows:
where:
P = pressure (p.s.i.);
C_{4} = 0.078 C_{1} (with C_{1} computed under 523.527);
K_{2} = hull station weighing factor, determined in accordance with figure Iii of Appendix I of this chapter;
V_{SO} = seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and
b = angle of dead rise at appropriate station.

(2) The unsymmetrical pressure distribution consists of the pressures prescribed in paragraph (c)(1) of this section on one side of the hull or main float centreline and onehalf of that pressure on the other side of the hull or main float centreline, in accordance with figure Iiii of Appendix I of this chapter.

(3) These pressures are uniform and must be applied simultaneously over the entire hull or main float bottom. The loads obtained must be carried into the sidewall structure of the hull proper, but need not be transmitted in a fore and aft direction as shear and bending loads.

(Change 5234 (960901))
523.535 Auxiliary Float Loads

(a) General. Auxiliary floats and their attachments and supporting structures must be designed for the conditions prescribed in this section. In the cases specified in paragraphs (b) through (e) of this section, the prescribed water loads may be distributed over the float bottom to avoid excessive local loads, using bottom pressures not less than those prescribed in paragraph (g) of this section.

(b) Step loading. The resultant water load must be applied in the plane of symmetry of the float at a point threefourths of the distance from the bow to the step and must be perpendicular to the keel. The resultant limit load is computed as follows, except that the value of L need not exceed three times the weight of the displaced water when the float is completely submerged:
where:
L = limit load (lbs.);
C_{5} = 0.0053;
V_{SO} = seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect;
W = seaplane design landing weight in pounds;
ß_{s} = angle of dead rise at a station 3/4 of the distance from the bow to the step, but need not be less than 15 degrees; and
r_{y} = ratio of the lateral distance between the centre of gravity and the plane of symmetry of the float to the radius of gyration in roll.

(c) Bow loading. The resultant limit load must be applied in the plane of symmetry of the float at a point onefourth of the distance from the bow to the step and must be perpendicular to the tangent to the keel line at that point. The magnitude of the resultant load is that specified in paragraph (b) of this section.

(d) Unsymmetrical step loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (a) of this section and a side component equal to 3.25 tan ß times the load specified in paragraph (b) of this section. The side load must be applied perpendicularly to the plane of symmetry of the float at a point midway between the keel and the chine.

(e) Unsymmetrical bow loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (b) of this section and a side component equal to 0.25 tan ß times the load specified in paragraph (c) of this section. The side load must be applied perpendicularly to the plane of symmetry at a point midway between the keel and the chine.

(f) Immersed float condition. The resultant load must be applied at the centroid of the cross section of the float at a point onethird of the distance from the bow to the step. The limit load components are as follows:
where:
? = mass density of water (slugs/ft.^{3})
V = volume of float(ft.^{3});
C_{X} = coefficient of drag force, equal to 0.133;
C_{Y} = coefficient of side force, equal to 0.l06;
K = 0.8, except that lower values may be used if it is shown that the floats are incapable of submerging at a speed of 0.8 V_{SO} in normal operations;
V_{SO} = seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and
g = acceleration due to gravity (ft/sec^{2}).

(g) Float bottom pressures. The float bottom pressures must be established under 523.533, except that the value of K_{2} in the formulae may be taken as 1.0. The angle of dead rise to be used in determining the float bottom pressures is set forth in paragraph (b) of this section.
(Change 5234 (960901))
523.537 Seawing Loads
Seawing design loads must be based on applicable test data.
(Change 5234 (960901))
Emergency Landing Conditions
523.561 General

(a) The aeroplane, although it may be damaged in emergency landing conditions, must be designed as prescribed in this section to 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 or 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 aeroplane, or 4.5g for aerobatic category aeroplanes;

(ii) Forward, 9.0 g;

(iii) Sideward, 1.5g; and

(iv) Downward, 6.0g when certification to the emergency exit provisions of 523.807(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, 18.0 g; 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 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 523.787(c), the supporting structure must be designed to restrain, under loads up to those specified in paragraph (b)(3) of this section, each item of mass that could injure an occupant if it came loose in a minor crash landing.

(1) For engines mounted inside the fuselage, aft of the cabin, it must be shown by test or analysis that the engine and attached accessories, and the engine mounting structure:
(effective 2016/08/04)
(i) can withstand a forward acting static ultimate inertia load factor of 18.0 g plus the maximum takeoff engine thrust; or

(ii) the aeroplane structure is designed to preclude the engine and its attached accessories from entering or protruding into the cabin should the engine mounts fail.


(2) (Reserved)
(effective 2016/08/04)

(Change 5231 (880101))
(Change 5232 (890101))
(Change 5235)
523.562 Emergency Landing Dynamic Conditions

(a) Each seat/restraint system for use in a normal, utility, or aerobatic category aeroplane, or in a commuter category jet aeroplane, must be designed to protect each occupant during an emergency landing when:
(effective 2016/08/04)
(1) Proper used is made of seats, safety belts, and shoulder harnesses provided for in the design; and

(2) The occupant is exposed to the loads resulting from the conditions prescribed in this section.


(b) Except for those seat/restraint systems that are required to meet paragraph (d) of this section, each seat/restraint system, for crew or passenger occupancy in a normal, utility, or aerobatic category aeroplane, or in a commuter category jet aeroplane, must successfully complete dynamic tests or be demonstrated by rational analysis supported by dynamic tests, in accordance with each of the following conditions. These tests must be conducted with an occupant simulated by an anthropomorphic test dummy (ATD) defined by Title 49 of the United States Code of Federal Regulations, Part 572, Subpart B or an equivalent document approved by the Minister, with a nominal weight of 170 pounds and seated in the normal upright position.
(effective 2016/08/04)
(1) For the first test, the change in velocity may not be less than 31 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the aeroplane and with the horizontal plane of the aeroplane pitched up 60 degrees, with no yaw, relative to the impact vector. For seat/restraint systems to be installed in the first row of the aeroplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 19g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 15g.

(2) For the second test, the change in velocity may not be less than 42 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the aeroplane and with the vertical plane of the aeroplane yawed 10 degrees, with no pitch, relative to the impact vector in a direction that results in the greatest load on the shoulder harness. For seat/restraint systems to be installed in the first row of the aeroplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 26g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 21g.

(3) To account for floor warpage, the floor rails or attachment devices used to attach the seat/restraint system to the airframe structure must be preloaded to misalign with respect to each other by at least 10 degrees vertically (i.e., pitch out of parallel) and one of the rails or attachment devices must be preloaded to misalign by 10 degrees in roll prior to conducting the test defined by paragraph (b)(2) of this section.


(c) Compliance with the following requirements must be shown during the dynamic tests conducted in accordance with paragraph (b) of this section:

(1) The seat/restraint system must restraint the ATD although seat/restraint system components may experience deformation, elongation, displacement, or crushing intended as part of the design.

(2) The attachment between the seat/restraint system and the test fixture must remain intact, although the seat structure may have deformed.

(3) Each shoulder harness strap must remain on the ATD's shoulder during the impact.

(4) The safety belt must remain on the ATD's pelvis during the impact.

(5) The results of the dynamic tests must show that the occupant is protected from serious head injury.

(i) When contact with adjacent seats, structure, or other items in the cabin can occur, protection must be provided so that the head impact does not exceed a head injury criteria (HIC) of 1,000.

(ii) The value of HIC is defined as:
Where:
t_{1} is the initial integration time, expressed in seconds, t_{2} is the final integration time, expressed in seconds, and a(t) is the total acceleration vs. time curve for the head strike expressed as a multiple of g (units of gravity).
(effective 2016/08/04) 
(iii) Compliance with the HIC limit must be demonstrated by measuring the head impact during dynamic testing as prescribed in paragraphs (b)(1) and (b)(2) of this section or by a separate showing of compliance with the head injury criteria using test or analysis procedures.


(6) Loads in individual shoulder harness straps may not exceed 1,750 pounds. If dual straps are used for retaining the upper torso, the total strap loads may not exceed 2,000 pounds.

(7) The compression load measured between the pelvis and the lumbar spine of the ATD may not exceed 1,500 pounds.


(d) For all singleengine aeroplanes with a V_{SO} of more than 61 knots at maximum weight, and those multiengine aeroplanes of 6,000 pounds or less maximum weight with a V_{SO} of more than 61 knots at maximum weight that do not comply with 523.67(a)(1);

(1) The ultimate load factors of 523.561(b) must be increased by multiplying the load factors by the square of the ratio of the increased stall speed to 61 knots. The increased ultimate load factors need not exceed the values reached at a V_{SO} of 79 knots. The upward ultimate load factor for aerobatic category aeroplanes need not exceed 5.0g.

(2) The seat/restraint system test required by paragraph (b)(1) of this section must be conducted in accordance with the following criteria:

(i) The change in velocity may not be less than 31 feet per second.

(ii)

(A) The peak deceleration (g_{p}) of 19g and 15g must be increased and multiplied by the square of the ratio of the increased stall speed to 61 knots:
g_{p} = 19.0(V_{so}/61)^{2} or
g_{p} = 15.0(V_{so}/61)^{2}

(B) The peak deceleration need not exceed the value reached at a V_{SO} of 79 knots.


(iii) The peak deceleration must occur in not more than time (t_{r),} which must be computed as follows:
where:
g _{p}= The peak deceleration calculated in accordance with paragraph (d)(2)(ii) of this section
t_{r} = The rise time (in seconds) to the peak deceleration.



(e) An alternate approach that achieves an equivalent, or greater, level of occupant protection to that required by this section may be used if substantiated on a rational basis.
(Change 5232 (890101))
(Change 5234 (960901))
(Change 5235)
Fatigue Evaluation
523.571 Metallic Pressurized Cabin 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;

(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 V_{C}, 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 effect of failure under static load are otherwise considered;

(c) The damage tolerance evaluation of 523.573(b); or

(d) If certification for operation above 41,000 feet is requested, a damage tolerance evaluation of the fuselage pressure boundary in accordance with 523.573(b) must be conducted.
(effective 2016/08/04)
(Change 5234 (960901))
(Change 5235)
523.572 Metallic Wing, Empennage, And Associated Structures

(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 at V_{C}. 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 523.573(b).


(b) Each evaluation required by this section must:

(1) Include typical loading spectra (e.g. taxi, groundairground 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.

(Change 5231 (880101))
(Change 5233 (920102))
(Change 5234 (960901))
(Change 5235)
523.573 Damage Tolerance and Fatigue Evaluation of Structure

(a) Composite airframe structure. Composite airframe structure must be evaluated under this paragraph instead of 523.571 and 523.572. The applicant must evaluate the composite airframe structure, the failure of which would result in catastrophic loss of the aeroplane, in each wing (including canards, tandem wings, and winglets), empennage, their carrythrough and attaching structure, moveable control surfaces and their attaching structure, fuselage, and pressure cabin using the damagetolerance criteria prescribed in paragraphs (a)(1) through (a)(4) of this section unless shown to be impractical. If the applicant establishes that damagetolerance criteria is impractical for a particular structure, the structure must be evaluated in accordance with paragraphs (a)(1) and (a)(6) of this section. Where bonded joints are used, the structure must also be evaluated in accordance with paragraph (a)(5) of this section. The effects of material variability and environmental conditions on the strength and durability properties of the composite materials must be accounted for in the evaluations required by this section.

(1) It must be demonstrated by tests, or by analysis supported by tests, that the structure is capable of carrying ultimate load with damage up to the threshold of detectability considering the inspection procedures employed.

(2) The growth rate or nogrowth of damage that may occur from fatigue, corrosion, manufacturing flaws or impact damage, under repeated loads expected in service, must be established by tests or analysis supported by tests.

(3) The structure must be shown by residual strength tests, or analysis supported by residual strength tests, to be able to withstand critical limit flight loads, considered as ultimate loads, with the extent of detectable damage consistent with the results of the damage tolerance evaluations. For pressurised cabins, the following loads must be withstood:

(i) Critical limit flight loads with the combined effects of normal operating pressure and expected external aerodynamic pressures.

(ii) The expected external aerodynamic pressures in 1g flight combined with a cabin differential pressure equal to 1.1 times the normal operating differential pressure without any other load.


(4) The damage growth, between initial detectability and the value selected for residual strength demonstrations, factored to obtain inspection intervals, must allow development of an inspection program suitable for application by operation and maintenance personnel.

(5) For any bonded joint, the failure of which would result in catastrophic loss of the aeroplane, the limit load capacity must be substantiated by one of the following methods:

(i) The maximum disbonds of each bonded joint consistent with the capability to withstand the loads in paragraph (a)(3) of this section must be determined by analysis, tests, or both. Disbonds of each bonded joint greater than this must be prevented by design features; or

(ii) Proof testing must be conducted on each production article that will apply the critical limit design load to each critical bonded joint; or

(iii) Repeatable and reliable nondestructive inspection techniques must be established that ensure the strength of each joint.


(6) Structural components for which the damage tolerance method is shown to be impractical must be shown by component fatigue tests, or analysis supported by tests, to be able to withstand the repeated loads of variable magnitude expected in service. Sufficient component, subcomponent, element, or coupon tests must be done to establish the fatigue scatter factor and the environmental effects. Damage up to the threshold of detectability and ultimate load residual strength capability must be considered in the demonstration.


(b) Metallic airframe structure.If the applicant elects to use 523.571(c) or 523.572(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:
(amended 2010/01/29; previous version )
(1) The normal operating differential pressure combined with the expected external aerodynamic pressures applied simultaneously with the flight loading conditions specified in this chapter, and

(2) The expected external aerodynamic pressures in 1g flight combined with a cabin differential pressure equal to 1.1 times the normal operating differential pressure without any other load.


(c) Removed
(Change 5234 (960901))
(Change 5235)
523.574 Metallic Damage Tolerance And Fatigue Evaluation Of Commuter Category Aeroplanes
For commuter category aeroplanes:

(a) Metallic damage tolerance. An evaluation of the strength, detail design, and fabrication must show that catastrophic failure due to fatigue, corrosion, defects, or damage will be avoided throughout the operational life of the aeroplane. This evaluation must be conducted in accordance with the provisions of 523.573, except as specified in paragraph (b) of this section, for each part of the structure that could contribute to a catastrophic failure.

(b) Fatigue (safelife) evaluation. Compliance with the damage tolerance requirements of paragraph (a) of this section is not required if the applicant establishes that the application of those requirements is impractical for a particular structure. This structure must be shown, by analysis supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected during its service life without detectable cracks. Appropriate safelife scatter factors must be applied.
(Change 5235)
523.575 Inspections And Other Procedures
Each inspection or other procedure, based on an evaluation required by 523.571, 523.572, 523.573 or 523.574, must be established to prevent catastrophic failure and must be included in the Limitations Section of the Instructions for Continued Airworthiness required by 523.1529.
(Change 5235)
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