Part V  Airworthiness Manual Chapter 525  Transport Category Aeroplanes
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Canadian Aviation Regulations (CARs) 20172
Preamble
SUBCHAPTERS
 A (525.1525.2),
 B (525.21525.255),
 C (525.301525.581),
 D (525.601525.899),
 E (525.901525.1207),
 F (525.1301525.1461),
 G (525.1501525.1587)
 H (525.1701525.1733)
APPENDICES
A, B, C, D, E, F, G, H, I, J, L, M, N
(2001/06/01; no previous version)
SUBCHAPTER C STRUCTURE  GENERAL
525.301 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 specified 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 must be validated by flight load measurement unless the methods used for determining those loading conditions are shown to be reliable.
 (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account.
525.303 Factor of Safety
Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit load which are considered external loads on the structure. When a loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified.
525.305 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 3 seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the 3second limit does not apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading. When analytical methods are used to show compliance with the ultimate load strength requirements, it must be shown that:
 (1) The effects of deformation are not significant;
 (2) The deformations involved are fully accounted for in the analysis; or
 (3) The methods and assumptions used are sufficient to cover the effects of these deformations.
 (c) Where structural flexibility is such that any rate of load application likely to occur in the operating conditions might produce transient stresses appreciably higher than those corresponding to static loads, the effects of this rate of application must be considered.
 (d) Removed and Reserved
 (e) The aeroplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to V_{D}/M_{D}, including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope. This must be shown by analysis, flight tests, or other tests found necessary by the Minister.
 (f) Unless shown to be extremely improbable, the aeroplane must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These must be considered limit loads and must be investigated at airspeeds up to V_{C}/M_{C}.
(Change 5255 (921030))
(Change 5258)
525.307 Proof of Structure
 (a) Compliance with the strength and deformation requirements of this subchapter must be shown for each critical loading condition. Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable. The Minister may require ultimate load tests in cases where limit load tests may be inadequate.
 (b) (Reserved)
 (c) (Reserved)
 (d) When static or dynamic tests are used to show compliance with the requirements of 525.305(b) for flight structures, appropriate material correction factors must be applied to the test results, unless the structure, or part thereof, being tested has features such that a number of elements contribute to the total strength of the structure and the failure of one element results in the redistribution of the load through alternate load paths.
(Change 5253 (911101)
Flight Loads
525.321 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 load factor is one in which the aerodynamic force acts upward with respect to the aeroplane.

(b) Considering compressibility effects at each speed, compliance with the flight load requirements of this subchapter must be shown:
 (1) At each critical altitude within the range of altitudes selected by the applicant;
 (2) At each weight from the design minimum weight to the design maximum weight appropriate to each particular flight load condition; and
 (3) For each required altitude and weight, for any practicable distribution of disposable load within the operating limitations recorded in the Aeroplane Flight Manual.
 (c) Enough points on and within the boundaries of the design envelope must be investigated to ensure that the maximum load for each part of the aeroplane structure is obtained.
 (d) The significant forces acting on the aeroplane must be placed in equilibrium in a rational or conservative manner. The linear inertia forces must be considered in equilibrium with the thrust and all aerodynamic loads, while the angular (pitching) inertia forces must be considered in equilibrium with thrust and all aerodynamic moments, including moments due to loads on components such as tail surfaces and nacelles. Critical thrust values in the range from zero to maximum continuous thrust must be considered.
(Change 5258)
Flight Manoeuvre and Gust Conditions
525.331 Symmetric Manoeuvring Conditions

(a) Procedure. For the analysis of the manoeuvring flight conditions specified in paragraphs (b) and (c) of this section, the following provisions apply:
 (1) Where sudden displacement of a control is specified, the assumed rate of control surface displacement may not be less than the rate that could be applied by the pilot through the control system.
 (2) In determining elevator angles and chordwise load distribution in the manoeuvring conditions of paragraphs (b) and (c) of this section, the effect of corresponding pitching velocities must be taken into account. The intrim and outoftrim flight conditions specified in 525.255 must be considered.
 (b) Manoeuvring balanced conditions. Assuming the aeroplane to be in equilibrium with zero pitching acceleration, the manoeuvring conditions A through I on the manoeuvring envelope in 525.333(b) must be investigated.

(c) Manoeuvring Pitching Conditions.The following conditions must be investigated:
(effective 2017/06/19) (1) Maximum pitch control displacement at V_{A}. The aeroplane is assumed to be flying in steady level flight (point A_{1}, 525.333(b)) and the cockpit pitch control is suddenly moved to obtain extreme nose up pitching acceleration. In defining the tail load, the response of the aeroplane must be taken into account. Aeroplane loads that occur subsequent to the time when normal acceleration at the c.g. exceeds the positive limit manoeuvring load factor (at point A_{2}, 525.333(b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.

(2) Checked manoeuvre between V_{A} and V_{D}. Noseup checked pitching manoeuvres must be analysed when the positive limit load factor prescribed in section 525.337 is achieved. As a separate condition, nosedown checked pitching manoeuvres in which a limit load factor of 0g is achieved must be analysed. In defining the aeroplane loads, the flight deck pitch control motions described in paragraphs (c)(2)(i) through (iv) of this section must be used:
(effective 2017/06/19)
(i) The aeroplane is assumed to be flying in steady level flight at any speed between V_{A} and V_{D} and the flight deck pitch control is moved in accordance with the following formula:
(effective 2017/06/19)δ(t) = δ_{1} sin(ωt) for 0 ≤ t ≤ t _{max}
Where
δ_{1} = the maximum available displacement of the flight deck pitch control in the initial direction, as limited by the control system stops, control surface stops, or by pilot effort in accordance with 525.397(b);
δ(t) = the displacement of the flight deck pitch control as a function of time. In the initial direction, δ(t) is limited to . In the reverse direction, δ(t) may be truncated at the maximum available displacement of the flight deck pitch control as limited by the control system stops, control surface stops, or by pilot effort in accordance with 525.397(b);
${\text{t}}_{\text{max}}\text{=3\pi /2\omega}$
ω = the circular frequency (radians/second) of the control deflection taken equal to the undamped natural frequency of the short period rigid mode of the aeroplane, with active control system effects included where appropriate; but not less than:
$\omega =\frac{\pi V}{\text{2}{\mathit{\text{V}}}_{\mathit{\text{A}}}}\text{radianspersecond;}$
Where
V = the speed of the aeroplane at entry to the manoeuvre;
V_{A} = the design manoeuvring speed prescribed in 525.335(c).

(ii) For noseup pitching manoeuvres, the complete flight deck pitch control displacement history may be scaled down in amplitude to the extent necessary to ensure that the positive limit load factor prescribed in section 525.337 is not exceeded. For nosedown pitching manoeuvres, the complete flight deck control displacement history may be scaled down in amplitude to the extent necessary to ensure that the normal acceleration at the centre of gravity does not go below 0g.
(effective 2017/06/19)  (iii) In addition, for cases where the aeroplane response to the specified flight deck pitch control motion does not achieve the prescribed limit load factors, then the following flight deck pitch control motion must be used:
(effective 2017/06/19)δ(t) = δ_{1} sin(ωt) for 0 ≤ t ≤ t _{1}
δ(t) = δ_{1} for t _{1}≤ t ≤ t _{2}
δ(t) = δ_{1} sin(ω[t + t _{1}− t _{2}]) for t _{2}≤ t ≤ t _{max}
Where
t _{1} = 𝜋/2ω
t _{2} = t _{1}+ Δt
t _{max} = t _{2}+ 𝜋/ω;
Δt = the minimum period of time necessary to allow the prescribed limit load factor to be achieved in the initial direction, but it need not exceed five seconds (see figure below).

(iv) In cases where the flight deck pitch control motion may be affected by inputs from systems (for example, by a stick pusher that can operate at high load factor as well as at 1g, then the effects of those systems shall be taken into account.
(effective 2017/06/19) 
(v) Aeroplane loads that occur beyond the following times need not be considered:
(effective 2017/06/19) (A) for the noseup pitching manoeuvre, the time at which the normal acceleration at the centre of gravity goes below 0g;
 (B) for the nosedown pitching manoeuvre, the time at which the normal acceleration at the centre of gravity goes above the positive limit load factor prescribed in section 525.337; or
 (C) t _{max}.

(i) The aeroplane is assumed to be flying in steady level flight at any speed between V_{A} and V_{D} and the flight deck pitch control is moved in accordance with the following formula:
 (d) Removed
525.333 Flight Manoeuvring Envelope
 (a) General. The strength requirements must be met at each combination of airspeed and load factor on and within the boundaries of the representative manoeuvring envelope (V_{n} diagram) of paragraph (b) of this section. This envelope must also be used in determining the aeroplane structural operating limitations as specified in 525.1501.
 (b) Manoeuvring Envelope.
 (c) Removed
(Change 5258)
525.335 Design Airspeeds
The selected design airspeeds are equivalent airspeeds (EAS). Estimated values of V_{SO} and V_{S1} must be conservative.

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

(1) The minimum value of V_{C} must be sufficiently greater than V_{B} to provide for inadvertent speed increases likely to occur as a result of severe atmospheric turbulence.
(amended 2012/03/27; previous version)  (2) Except as provided in 525.335(d)(2), V_{C} may not be less than V_{B} + 1.32 U_{REF} (with U_{REF} as specified in 525.341(a)(5)(i)). However, V_{C} need not exceed the maximum speed in level flight at maximum continuous power for the corresponding altitude.
 (3) At altitudes where V_{D} is limited by Mach number, V_{C} may be limited to a selected Mach number.

(1) The minimum value of V_{C} must be sufficiently greater than V_{B} to provide for inadvertent speed increases likely to occur as a result of severe atmospheric turbulence.

(b) Design dive speed, V_{D}. V_{D} must be selected so that V_{C}/M_{C} is not greater than 0.8 V_{D}/M_{D}, or so that the minimum speed margin between V_{C}/M_{C} and V_{D}/M_{D} is the greater of the following values:
 (1) From an initial condition of stabilised flight at V_{C}/M_{C}, the aeroplane is upset, flown for 20 seconds along a flight path of 7.5° below the initial path, and then pulled up at a load factor of 1.5g (0.5g acceleration increment). The speed increase occurring in this manoeuvre may be calculated if reliable or conservative aerodynamic data issued. Power as specified in 525.175(b)(1)(iv) is assumed until the pullup is initiated, at which time power reduction and the use of pilot controlled drag devices may be assumed;
 (2) The minimum speed margin must be enough to provide for atmospheric variations (such as horizontal gusts, and penetration of jet streams and cold fronts) and for instrument errors and airframe production variations. These factors may be considered on a probability basis. The margin at altitude where M_{C} is limited by compressibility effects must not be less than 0.07M unless a lower margin is determined using a rational analysis that includes the effects of any automatic systems. In any case, the margin may not be reduced to less than 0.05M.

(c) Design manoeuvring speed, V_{A}. For V_{A}, the following apply:
 (1) V_{A} may not be less than
${V}_{S1}\sqrt{\mathit{\text{n}}}$
where:


 (i) n is the limit positive manoeuvring load factor at V_{C}; and
 (ii) V_{S1} is the stalling speed with flaps retracted.
 (2) V_{A} and V_{S} must be evaluated at the design weight and altitude under consideration.
 (3) V_{A} need not be more than V_{C} or the speed at which the positive C_{Nmax} curve intersects the positive manoeuvre load factor line, whichever is less.


(d) Design speed for maximum gust intensity, V_{B}.
 (1) V_{B} may not be less than:
${V}_{\mathit{S}\mathit{\text{1}}}{\left[\mathit{1}+\frac{{\mathit{\text{K}}}_{\mathit{g}}{\mathit{\text{U}}}_{\mathit{\text{ref}}\mathit{}}{\mathit{\text{V}}}_{\mathit{c}}\mathit{\text{a}}}{\text{498w}}\right]}^{\raisebox{1ex}{$1$}\!\left/ \!\raisebox{1ex}{$2$}\right.}$
where:
V_{S1} = the 1g stalling speed based on C_{NAmax} with the flaps retracted at the particular weight under consideration;
V_{c} = design cruise speed (knots equivalent airspeed);
U_{ref} = the reference gust velocity (feet per second equivalent airspeed) from 525.341(a)(5)(i);
w = average wing loading (pounds per square foot) at the particular weight under consideration.
${\mathit{\text{K}}}_{\mathit{g}}\text{=}\frac{\text{0,88}\mu}{\text{5,3+}\mathit{\text{\mu}}}$
$\mu =\frac{2\mathrm{w}}{\mathit{\text{pcag}}}$
r = density of air (slugs/ft^{3});
c = mean geometric chord of the wing (feet);
g = acceleration due to gravity (ft/sec^{2});
a = slope of the aeroplane normal force coefficient curve, C_{NA} per radian;


(2) At altitudes where V_{C} is limited by Mach number:
 (i) V_{B} may be chosen to provide an optimum margin between low and high speed buffet boundaries; and,
 (ii) V_{B} need not be greater than V_{c}.

(2) At altitudes where V_{C} is limited by Mach number:

(e) Design flap speeds, V_{F}. For V_{F}, the following apply:
 (1) The design flap speed for each flap position (established in accordance with 525.697(a)) must be sufficiently greater than the operating speed recommended for the corresponding stage of flight (including balked landings) to allow for probable variations in control of airspeed and for transition from one flap position to another.
 (2) If an automatic flap positioning or load limiting device is used, the speeds and corresponding flap positions programmed or allowed by the device may be used.

(3) V_{F} may not be less than:
 (i) 1.6 V_{S1} with the flaps in takeoff position at maximum takeoff weight;
 (ii) 1.8 V_{S1} with the flaps in approach position at maximum landing weight; and
 (iii) 1.8 V_{SO} with the flaps in landing position at maximum landing weight.
 (f) Design drag device speeds, V_{DD}. The selected design speed for each drag device must be sufficiently greater than the speed recommended for the operation of the device to allow for probable variations in speed control. For drag devices intended for use in high speed descents, V_{DD} may not be less than V_{D}. When an automatic drag device positioning or load limiting means is used, the speeds and corresponding drag device positions programmed or allowed by the automatic means must be used for design.
(Change 5258)
525.337 Limit Manoeuvring Load Factors
 (a) Except where limited by maximum (static) lift coefficients, the aeroplane is assumed to be subjected to symmetrical manoeuvres resulting in the limit manoeuvring load factors prescribed in this section. Pitching velocities appropriate to the corresponding pullup and steady turn manoeuvres must be taken into account.
 (b) The positive limit manoeuvring load factor "n" for any speed up to V_{N} may not be less than
$\text{2.1+}\left(\frac{\text{24,000}}{\mathit{\text{W}}\text{+10,000}}\right)$
except that "n" may not be less than 2.5 and need not be greater than 3.8 where "W" is the design maximum takeoff weight.

(c) The negative limit manoeuvring load factor:
 (1) May not be less than 1.0 at speeds up to V_{C}; and
 (2) Must vary linearly with speed from the value at V_{C} to zero at V_{D}.
 (d) 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 5253 (911101))
525.341 Gust and Turbulence Loads

(a) Discrete Gust Design Criteria. The aeroplane is assumed to be subjected to symmetrical vertical and lateral gusts in level flight. Limit gust loads must be determined in accordance with the provisions:
 (1) Loads on each part of the structure must be determined by dynamic analysis. The analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions.

(2) The shape of the gust must be:
$\mathit{\text{U}}\text{=}\frac{{\mathit{\text{U}}}_{\mathit{\text{ds}}}}{\mathit{\text{2}}}\left[\mathit{\text{1}}\text{}\mathrm{cos}\left(\frac{\pi s}{\mathit{\text{H}}}\right)\right]$
for 0 < s < 2H
where:
s = distance penetrated into the gust (feet);
U_{ds} = the design gust velocity in equivalent airspeed specified in paragraph (a)(4) of this section; and
H = the gust gradient which is the distance (feet) parallel to the aeroplane's flight path for the gust to reach its peak velocity.
 (3) A sufficient number of gust gradient distances in the range 30 feet to 350 feet must be investigated to find the critical response for each load quantity.

(4) The design gust velocity must be:
U_{ds}= U_{ref}F_{g}(^{H}/350)^{1/6}
where:
U_{ref} = the reference gust velocity in equivalent airspeed defined in paragraph (a)(5) of this section.
F_{g} = the flight profile alleviation factor defined in paragraph (a)(6) of this section.

(5) The following reference gust velocities apply:

(i) At aeroplane speeds between V_{B} and V_{C}:
(effective 2017/06/19) 
(ii) At the aeroplane design speed V_{D}:
The reference gust velocity must be 0.5 times the value obtained under 525.341(a)(5)(i).
Positive and negative gusts with reference gust velocities of 56.0 ft/sec EAS must be considered at sea level. The reference gust velocity may be reduced linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS at 15,000 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at 15,000 feet to 20.86 ft/sec EAS at 60,000 feet.
(effective 2017/06/19) 
(i) At aeroplane speeds between V_{B} and V_{C}:

(6) The flight profile alleviation factor, F_{g}, must be increased linearly from the sea level value to a value of 1.0 at the maximum operating altitude defined in 525.1527. At sea level, the flight profile alleviation factor is determined by the following equation:
${\mathit{\text{F}}}_{\mathit{\text{g}}}\text{=0,5}\left({F}_{\mathit{g}\mathit{z}}\text{+}{\mathit{\text{F}}}_{\mathit{g}\mathit{m}}\right)$
where:
${\mathit{\text{F}}}_{\mathit{g}\mathit{m}}\text{=1}\frac{{\mathit{\text{Z}}}_{\mathit{m}\mathit{o}}}{250000}$
${\mathit{\text{F}}}_{\mathit{g}\mathit{m}}\text{=}\sqrt{{\mathit{\text{R}}}_{\mathit{2}}\mathit{\text{Tan}}\left({}^{\mathit{\pi}\mathit{R}}\mathit{1}\mathit{/}\mathit{4}\right)}$
${\text{R}}_{1}\text{=}\frac{\text{MaximumLandingWeight}}{\text{MaximumTakeoffWeight}}\text{;}$
${\mathrm{R}}_{2}=\frac{\mathrm{Maximum}\mathrm{Zero}\mathrm{Fuel}\mathrm{Weight}}{\mathrm{Maximum}\mathrm{Take}\mathrm{off}\mathrm{Weight}};$
Z_{mo} = Maximum operating altitude defined in section 525.1527 (feet).
(effective 2017/06/19)  (7) When a stability augmentation system is included in the analysis, the effect of any significant system nonlinearities should be accounted for when deriving limit loads from limit gust conditions.

(b) Continuous Turbulence Design Criteria. The dynamic response of the aeroplane to vertical and lateral continuous turbulence must be taken into account. The dynamic analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions. The limit loads must be determined for all critical altitudes, weights, and weight distributions as specified in 525.321(b), and all critical speeds within the ranges indicated in 525.341(b)(3).
(effective 2017/06/19) (1) Except as provided in paragraphs (b)(4) and (5) of this section, the following equation must be used:
(effective 2017/06/19)${\mathrm{P}}_{\mathrm{L}}={\mathrm{P}}_{\mathrm{L}1\mathrm{g}}\pm {\mathrm{U}}_{\mathrm{\sigma}}\overline{\mathrm{A}}$
Where
P_{L} = limit load;
P_{L1g} = steady 1g load for the condition;
Ā = ratio of rootmeansquare incremental load for the condition to rootmeansquare turbulence velocity; and
U_{σ} = limit turbulence intensity in true airspeed, specified in paragraph (b)(3) of this section.
 (2) Values of Ā must be determined according to the following formula:
(effective 2017/06/19)$\overline{A}=\sqrt{{}_{0}{}^{\infty}\int {\leftH\left(\mathrm{\Omega}\right)\right}^{2}\mathrm{\Phi}\left(\mathrm{\Omega}\right)d\mathrm{\Omega}}$
Where
H(Ω) = the frequency response function, determined by dynamic analysis, that relates the loads in the aircraft structure to the atmospheric turbulence; and
Φ(Ω) = normalized power spectral density of atmospheric turbulence given by:
$\mathrm{\Phi}\left(\mathrm{\Omega}\right)=\frac{L}{\pi}\frac{1+{\displaystyle \frac{8}{3}}{\left(1.339\mathrm{\Omega}L\right)}^{2}}{{\left[1+{\left(1.339\mathrm{\Omega}L\right)}^{2}\right]}^{{\displaystyle \raisebox{1ex}{$11$}\!\left/ \!\raisebox{1ex}{$6$}\right.}}}$
Where
Ω = reduced frequency, radians per foot; and
L = scale of turbulence = 2,500 ft.
 (3) The limit turbulence intensities, Uσ, in feet per second true airspeed required for compliance with this paragraph are:
(effective 2017/06/19) (i) at aeroplane speeds between V_{B} and V_{C}: U_{σ} = U_{σref}F_{g}
Where
U_{σref} is the reference turbulence intensity that varies linearly with altitude from 90 fps (TAS) at sea level to 79 fps (TAS) at 24,000 feet and is then constant at 79 fps (TAS) up to the altitude of 60,000 feet;
F_{g} is the flight profile alleviation factor defined in paragraph (a)(6) of this section;
 (ii) at speed V_{D}: U_{σ} is equal to ½ the values obtained under paragraph (b)(3)(i) of this section
 (iii) at speeds between V_{C} and V_{D}: U_{σ} is equal to a value obtained by linear interpolation; and
 (iv) at all speeds, both positive and negative incremental loads due to continuous turbulence must be considered.
 (i) at aeroplane speeds between V_{B} and V_{C}: U_{σ} = U_{σref}F_{g}
 (4) When an automatic system affecting the dynamic response of the aeroplane is included in the analysis, the effects of system nonlinearities on loads at the limit load level must be taken into account in a realistic or conservative manner.
(effective 2017/06/19)  (5) If necessary for the assessment of loads on aeroplanes with significant nonlinearities, it must be assumed that the turbulence field has a rootmeansquare velocity equal to 40 percent of the U_{σ} values specified in paragraph (b)(3) of this section. The value of limit load is that load with the same probability of exceedance in the turbulence field as ĀU_{σ} of the same load quantity in a linear approximated model.
(effective 2017/06/19)
 (1) Except as provided in paragraphs (b)(4) and (5) of this section, the following equation must be used:
 (c) Supplementary gust conditions for wingmounted engines. For aeroplanes equipped with wingmounted engines, the engine mounts, pylons, and wing supporting structure must be designed for the maximum response at the nacelle centre of gravity derived from the following dynamic gust conditions applied to the aeroplane:
(effective 2017/06/19) (1) a discrete gust determined in accordance with 525.341(a) at each angle normal to the flight path, and separately;
 (2) a pair of discrete gusts, one vertical and one lateral. The length of each of these gusts must be independently tuned to the maximum response in accordance with 525.341(a). The penetration of the aeroplane in the combined gust field and the phasing of the vertical and lateral component gusts must be established to develop the maximum response to the gust pair. In the absence of a more rational analysis, the following formula must be used for each of the maximum engine loads in all six degrees of freedom:
${\mathit{\text{P}}}_{L}\text{=}{\mathit{\text{P}}}_{\mathit{L}\mathit{}\mathit{1}\mathit{g}}\pm 0.85\sqrt{{L}_{\mathrm{V}}^{2}+{L}_{\mathit{L}}^{\mathit{2}}}$
Where
P_{L} = limit load;
P_{L1g} = steady 1g load for the condition;
L_{V} = peak incremental response load due to a vertical gust according to 525.341(a); and
L_{L} = peak incremental response load due to a lateral gust according to 525.341(a).
525.343 Design Fuel and Oil Loads
 (a) The disposable load combinations must include each fuel and oil load in the range from zero fuel and oil to the selected maximum fuel and oil load. A structural reserve fuel condition, not exceeding 45 minutes of fuel under the operating conditions in 525.1001(e) and (f), as applicable, may be selected.

(b) 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 as prescribed in this subchapter. In addition:

(1) The structure must be designed for a condition of zero fuel and oil in the wing at limit loads corresponding to:
 (i) a manoeuvring load factor of +2.25; and

(ii) the gust and turbulence conditions of 525.341(a) and (b), but assuming 85% of the design velocities prescribed in 525.341(a)(4) and 85% of the turbulence intensities prescribed in 525.341(b)(3).
(effective 2017/06/19)
 (2) Fatigue evaluation of the structure must account for any increase in operating stresses resulting from the design condition of subparagraph (b)(1) of this paragraph; and
 (3) The flutter, deformation, and vibration requirements must also be met with zero fuel.

(1) The structure must be designed for a condition of zero fuel and oil in the wing at limit loads corresponding to:
(Change 5253 (911101))
(Change 5258)
525.345 High Lift Devices

(a) If wing flaps are to be used during takeoff, approach, or landing, at the design flap speeds established for these stages of flight under 525.335(e) and with the wing flaps in the corresponding positions, the aeroplane is assumed to be subjected to symmetrical manoeuvres and gusts. The resulting limit loads must correspond to the conditions determined as follows:
 (1) Manoeuvring to a positive limit load factor of 2.0; and

(2) Positive and negative gusts of 25 ft/sec EAS acting normal to the flight path in level flight. Gust loads resulting on each part of the structure must be determined by rational analysis. The analysis must take into account the unsteady aerodynamic characteristics and rigid body motions of the aircraft. The shape of the gust must be as described in 525.341(a)(2) except that:
U_{ds} = 25 ft/sec EAS;
H = 12.5 c; and
c = mean geometric chord of the wing (feet).

(b) The aeroplane must be designed for the conditions prescribed in paragraph (a) of this section, except that the aeroplane load factor need not exceed 1.0, taking into account, as separate conditions, the effects of:
 (1) Propeller slipstream corresponding to maximum continuous power at the design flap speeds V_{F}, and with takeoff power at not less than 1.4 times the stalling speed for the particular flap position and associated maximum weight; and
 (2) A headon gust of 25 feet per second velocity (EAS).

(c) If flaps or other high lift devices are to be used in en route conditions, and with flaps in the appropriate position at speeds up to the flap design speed chosen for these conditions, the aeroplane is assumed to be subjected to symmetrical manoeuvres and gusts within the range determined by:
 (1) manoeuvring to a positive limit load factor as prescribed in 525.337(b); and

(2) the vertical gust and turbulence conditions prescribed in 525.341(a) and (b).
(effective 2017/06/19)
 (d) The aeroplane must be designed for a manoeuvring load factor of 1.5 g at the maximum takeoff weight with the wingflaps and similar high lift devices in the landing configurations.
(Change 5253 (911101))
(Change 5258)
525.349 Rolling Conditions
The aeroplane must be designed for loads resulting from the rolling conditions specified in paragraphs (a) and (b) of this section. 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.

(a) Manoeuvring. The following conditions, speeds, and aileron deflections (except as the deflections may be limited by pilot effort) must be considered in combination with an aeroplane load factor of zero and of twothirds of the positive manoeuvring factor used in design. In determining the required aileron deflections, the torsional flexibility of the wing must be considered in accordance with 525.301(b):
 (1) Conditions corresponding to steady rolling velocities must be investigated. In addition, conditions corresponding to maximum angular acceleration must be investigated for aeroplanes with engines or other weight concentrations outboard of the fuselage. For the angular acceleration conditions, zero rolling velocity may be assumed in the absence of a rational time history investigation of the manoeuvre.
 (2) At V_{A}, a sudden deflection of the aileron to the stop is assumed.
 (3) At V_{C}, the aileron deflection must be that required to produce a rate of roll not less than that obtained in subparagraph (2) of this paragraph.
 (4) At V_{D}, the aileron deflection must be that required to produce a rate of roll not less than onethird of that in subparagraph (2) of this paragraph.
 (b) Unsymmetrical gusts. The aeroplane is assumed to be subjected to unsymmetrical vertical gusts in level flight. The resulting limit loads must be determined from either the wing maximum airload derived directly from 525.341(a), or the wing maximum airload derived indirectly from the vertical load factor calculated from 525.341(a). It must be assumed that 100 percent of the wing air load acts on one side of the aeroplane and 80 percent of the wing air load acts on the other side.
(Change 5258)
525.351 Yaw Manoeuvre Conditions
The aeroplane must be designed for loads resulting from the yaw manoeuvre conditions specified in paragraphs (a) through (d) of this section at speeds from V_{MC} to V_{D}. Unbalanced aerodynamic moments about the centre of gravity must be reacted in a rational or conservative manner considering the aeroplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero.

(a) With the aeroplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by:
 (1) The control system on control surface stops; or
 (2) A limit pilot force of 300 pounds from V_{MC} to V_{A} and 200 pounds from V_{C}/M_{C} to V_{D}/M_{D}, with a linear variation between V_{A} and V_{C}/M_{C}.
 (b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the aeroplane yaws to the overswing sideslip angle.
 (c) With the aeroplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section.
 (d) With the aeroplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral.
(Change 5253 (911101))
(Change 5258)
Supplementary Conditions
525.361 Engine and Auxiliary Power Unit Torque
(effective 2017/06/19)

(a) For engine installations:
(effective 2017/06/19) (1) Each engine mount, pylon and adjacent supporting airframe structures must be designed for the effects of:
(effective 2017/06/19)
(i) a limit engine torque corresponding to takeoff power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with 75 percent of the limit loads from flight condition A of 525.333(b);
(effective 2017/06/19) 
(ii) a limit engine torque corresponding to the maximum continuous power/thrust and, if applicable, corresponding propeller speed acting simultaneously with the limit loads from flight condition A of 525.333(b); and
(effective 2017/06/19) 
(iii) for turbopropeller installations only, in addition to the conditions specified in subparagraphs (a)(1)(i) and (ii) of this paragraph, a limit engine torque corresponding to take off 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.
(effective 2017/06/19)

(i) a limit engine torque corresponding to takeoff power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with 75 percent of the limit loads from flight condition A of 525.333(b);
 (2) The limit engine torque to be considered under paragraph (a)(1) of this section must:
(effective 2017/06/19) (i) for turbopropeller installations, be obtained by multiplying the mean torque for the specified power/thrust and speed by a factor of 1.25; or
(effective 2017/06/19)  (ii) for other turbine engines, be equal to the maximum accelerating torque for the case considered.
(effective 2017/06/19)
 (i) for turbopropeller installations, be obtained by multiplying the mean torque for the specified power/thrust and speed by a factor of 1.25; or
 (3) The engine mounts, pylons, and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit engine torque loads imposed by each of the following conditions to be considered separately:
(effective 2017/06/19) (i) sudden maximum engine deceleration due to malfunction or abnormal condition; and
(effective 2017/06/19)  (ii) the maximum acceleration of engine.
(effective 2017/06/19)
 (i) sudden maximum engine deceleration due to malfunction or abnormal condition; and
 (1) Each engine mount, pylon and adjacent supporting airframe structures must be designed for the effects of:

(b) For auxiliary power unit installations, the power unit mounts and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit torque loads imposed by each of the following conditions to be considered separately:
(effective 2017/06/19)
(1) sudden maximum auxiliary power unit deceleration due to malfunction, abnormal condition, or structural failure; and
(effective 2017/06/19) 
(2) the maximum acceleration of the auxiliary power unit.
(effective 2017/06/19)

(1) sudden maximum auxiliary power unit deceleration due to malfunction, abnormal condition, or structural failure; and
(Change 5253 (911101))
525.362 Engine Failure Loads
(effective 2017/06/19)
 (a) For engine mounts, pylons, and adjacent supporting airframe structure, an ultimate loading condition must be considered that combines 1g flight loads with the most critical transient dynamic loads and vibrations, as determined by dynamic analysis, resulting from failure of a blade, shaft, bearing or bearing support, or bird strike event. Any permanent deformation from these ultimate load conditions must not prevent continued safe flight and landing.
 (b) The ultimate loads developed from the conditions specified in paragraph (a) of this section are to be:
 (1) multiplied by a factor of 1.0 when applied to engine mounts and pylons; and
 (2) multiplied by a factor of 1.25 when applied to adjacent supporting airframe structure.
525.363 Side Load on Engine and Auxiliary Power Unit Mounts

(a) Each engine and auxiliary power unit mount and its supporting structure must be designed for a limit load factor in lateral direction, for the side load on the engine and auxiliary power unit mount, at least equal to the maximum load factor obtained in the yawing conditions but not less than:
 (1) 1.33; or
 (2) Onethird of the limit load factor for flight condition A as prescribed in 525.333(b).
 (b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions.
(Change 5258)
525.365 Pressurised Compartment Loads
For aeroplanes with one or more pressurised compartments, 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 stress concentrations and fatigue effects must be accounted for.
 (c) If landings may be made with the compartment 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 designed to be able to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33 for aeroplanes to be approved for operation to 45,000 feet or by a factor of 1.67 for aeroplanes to be approved for operation above 45,000 feet, omitting other loads.

(e) Any structure, component or part inside or outside a pressurised compartment, the failure of which could interfere with continued safe flight and landing, must be designed to withstand the effects of a sudden release of pressure through an opening in any compartment at any approved operating altitude resulting from each of the following conditions:
 (1) The penetration of the compartment by a portion of an engine following an engine disintegration;
 (2) Any opening in any pressurised compartment up to the size H_{o} in square feet; however, small compartments may be combined with an adjacent pressurised compartment and both considered as a single compartment for openings that cannot reasonably be expected to be confined to the small compartment. The size H_{o} must be computed by the following formula:
_{${\text{H}}_{\xb0}{\text{=PA}}_{\mathrm{S}}$}
where:
H_{o} = maximum opening in square feet, not to exceed 20 square feet;
P = A_{S}/6,240 + 0.024;
A_{S} = maximum cross sectional area of pressurised shell normal to the longitudinal axis, in square feet; and

 (3) The maximum opening caused by aeroplane or equipment failures not shown to be extremely improbable.
 (f) In complying with paragraph (e) of this section, the failsafe features of the design may be considered in determining the probability of failure or penetration and probable size of openings, provided that possible improper operation of closure devices and inadvertent door openings are also considered. Furthermore, the resulting differential pressure loads must be combined in a rational and conservative manner with 1g level flight loads and any loads arising from emergency depressurisation conditions. These loads may be considered as ultimate conditions; however, any deformations associated with these conditions must not interfere with continued safe flight and landing. The pressure relief provided by intercompartment venting may also be considered.
 (g) Bulkheads, floors and partitions in pressurised compartments for occupants must de designed to withstand the conditions specified in paragraph (e) of this section. In addition, reasonable design precautions must be taken to minimise the probability of parts becoming detached and injuring occupants while in their seats.
(Change 5253 (911101))
(Change 5258)
525.367 Unsymmetrical Loads Due to Engine Failure

(a) The aeroplane must be designed for the unsymmetrical loads resulting from the failure of the critical engine. Turbopropeller aeroplanes must be designed for 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 enginepropeller aeroplane combination.
 (b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than two seconds after the engine failure. The magnitude of the corrective action may be based on the control forces specified in 525.397(b) 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.
525.371 Gyroscopic Loads
The structure supporting any engine or auxiliary power unit must be designed for the loads, including gyroscopic loads, arising from the conditions specified in sections 525.331, 525.341, 525.349, 525.351, 525.473, 525.479, and 525.481, with the engine or auxiliary power unit at the maximum rotating speed appropriate to the condition. For the purposes of compliance with this section, the pitch manoeuvre in 525.331(c)(1) must be carried out until the positive limit manoeuvring load factor (point A_{2} in 525.333(b)) is reached.
(effective 2017/06/19)
(Change 5258)
525.373 Speed Control Devices
If speed control devices (such as spoilers and drag flaps) are installed for use in en route conditions:

(a) The aeroplane must be designed for the symmetrical manoeuvres prescribed in sections 525.333 and 525.337, the yawing manoeuvres in section 525.351 and the vertical and lateral gust and turbulence conditions prescribed in 525.341(a) and (b), at each setting and the maximum speed associated with that setting; and
(effective 2017/06/19)  (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.
(Change 5253 (911101))
(Change 5258)
Control Surface and System Loads
525.391 Control Surface Loads: General
The control surfaces must be designed for the limit loads resulting from the flight conditions in section 525.331, 525.341(a) and (b), sections 525.349 and 525.351, considering the requirements for:
(effective 2017/06/19)
 (a) Loads parallel to hinge line, in 525.393;
 (b) Pilot effort effects, in 525.397;
 (c) Trim tab effects, in 525.407;
 (d) Unsymmetrical loads, in 525.427; and
 (e) Auxiliary aerodynamic surfaces, in 525.445.
(Change 5258)
525.393 Loads Parallel to Hinge Line
 (a) Control surfaces and supporting hinge brackets must be designed for inertia loads acting parallel to the hinge line.

(b) In the absence of more rational data, the inertia 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.
525.395 Control System
 (a) Longitudinal, lateral, directional and drag control systems and their supporting structures must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in 525.391.

(b) The system limit loads of paragraph (a) of this section need not exceed the loads that can be produced by the pilot (or pilots) and by automatic or power devices operating the controls.
(effective 2017/06/19)  (c) The loads must not be less than those resulting from application of the minimum forces prescribed in 525.397(c).
(Change 5253 (911101))
525.397 Control System Loads
 (a) General. The maximum and minimum pilot forces, specified in paragraph (c) of this section, are assumed to act at the appropriate control grips or pads (in a manner simulating flight conditions) and to be reacted at the attachment of the control system to the control surface horn.
 (b) Pilot effort effects. 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 (c) of this section. Two thirds of the maximum values specified for the aileron and elevator may be used if control surface hinge moments are based on reliable data. In applying this criterion, the effects of servo mechanisms, tabs, and automatic pilot systems, must be considered.
 (c) Limit pilot forces and torques. The limit pilot forces and torques are as follows:
Control  Maximum forces or torques  Minimum forces or torques 

Aileron:  
Stick  100 lbs.  40 lbs. 
Wheel^{1}  80 D in.lbs.^{2}  40 D in.lbs. 
Elevator:  
Stick  250 lbs.  100 lbs. 
Wheel (symmetrical)  300 lbs.  100 lbs. 
Wheel (unsymmetrical)^{3}  100 lbs.  
Rudder  300 lbs.  130 lbs. 
^{1} The critical parts of the aileron control system must be designed for a single tangential force with a limit value equal to 1.25 times the couple force determined from these criteria. ^{2} D = wheel diameter (inches) ^{3} The unsymmetrical forces must be applied at one of the normal handgrip points on the periphery of the control wheel. 
525.399 Dual Control System

(a) Each dual control system must be designed for the pilots operating in opposition, using individual pilot forces not less than:
 (1) 0.75 times those obtained under 525.395; or
 (2) The minimum forces specified in 525.397 (c).
 (b) The control system must be designed for pilot forces applied in the same direction, using individual pilot forces not less than 0.75 times those obtained under 525.395.
525.405 Secondary Control System
Secondary controls, such as wheel brake, spoiler, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls. The following values may be used:
Pilot Control Force Limits (Secondary Controls)
Control  Limit Pilot Forces 

Miscellaneous: *Crank, wheel, or lever. 
$\frac{1+R}{3}$ x 50 lbs., but not less than 50 lbs. nor more than 150 lbs. (R=radius) (Applicable to any angle within 20° of plane of control). 
Twist  133 in.lbs. 
Pushpull  To be chosen by applicant. 
* Limited to flap, tab, stabiliser, spoiler, and landing gear operation controls. 
525.407 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, and the deflections are:
 (a) For elevator trim tabs, those required to trim the aeroplane at any point within the positive portion of the pertinent flight envelope in 525.333(b), except as limited by the stops; and
 (b) For aileron and rudder trim tabs, those required to trim the aeroplane in the critical unsymmetrical power and loading conditions, with appropriate allowance for rigging tolerances.
525.409 Tabs
 (a) Trim tabs. Trim tabs must be designed to withstand loads arising from all likely combinations of tab setting, primary control position, and aeroplane speed (obtainable without exceeding the flight load conditions prescribed for the aeroplane as a whole), when the effect of the tab is opposed by pilot effort forces up to those specified in 525.397(b).
 (b) Balancing tabs. Balancing tabs must be designed for deflections consistent with the primary control surface loading conditions.
 (c) Servo tabs. Servo tabs must be designed for deflections consistent with the primary control surface loading conditions obtainable within the pilot manoeuvring effort, considering possible opposition from the trim tabs.
525.415 Ground Gust Conditions

(a) The flight control systems and surfaces must be designed for the limit loads generated when the aeroplane is subjected to a horizontal 65knot ground gust from any direction while taxiing and while parked. For aeroplanes equipped with control system gust locks, the taxiing condition must be evaluated with the controls locked and unlocked, and the parked condition must be evaluated with the controls locked.
(effective 2017/06/19) 
(b) The control system and surface loads due to ground gust may be assumed to be static loads, and the hinge moments H must be computed from the formula:
(effective 2017/06/19)$\text{H=K}\left(1/2\right)\text{}{\mathrm{\rho}}_{o}{\mathrm{V}}^{2}\mathrm{c}\mathrm{S}$
Where
K = hinge moment factor for ground gusts derived in paragraph (c) of this section;
(effective 2017/06/19)${\mathrm{\rho}}_{o}$ = density of air at sea level;
(effective 2017/06/19)V = 65 knots relative to the aircraft;
(effective 2017/06/19)S = area of the control surface aft of the hinge line;
(effective 2017/06/19)c = mean aerodynamic chord of the control surface aft of the hinge line.
(effective 2017/06/19)  (c) The hinge moment factor K for ground gusts must be taken from the following table:
(effective 2017/06/19)Surface K Position Of Controls (1) Aileron 0.75 Control column locked or lashed in midposition. (2) Aileron *±0.50 Ailerons at full throw. (3) Elevator *±0.75 Elevator full down. (4) Elevator *±0.75 Elevator full up. (5) Rudder 0.75 Rudder in neutral. (6) Rudder 0.75 Rudder at full throw. *A positive value of K indicates a moment tending to depress the surface, while a negative value of K indicates a moment tending to raise the surface.  (d) The computed hinge moment of paragraph (b) of this section must be used to determine the limit loads due to ground gust conditions for the control surface. A 1.25 factor on the computed hinge moments must be used in calculating limit control system loads.
(effective 2017/06/19)  (e) Where control system flexibility is such that the rate of load application in the ground gust conditions might produce transient stresses appreciably higher than those corresponding to static loads, in the absence of a rational analysis substantiating a different dynamic factor, an additional factor of 1.6 must be applied to the control system loads of paragraph (d) of this section to obtain limit loads. If a rational analysis is used, the additional factor must not be less than 1.2.
(effective 2017/06/19)  (f) For the condition of the control locks engaged, the control surfaces, the control system locks, and the parts of any control systems between the surfaces and the locks must be designed to the resultant limit loads. Where control locks are not provided, then the control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads. If the control system design is such as to allow any part of the control system to impact with the stops due to flexibility, then the resultant impact loads must be taken into account in deriving the limit loads due to ground gust.
(effective 2017/06/19)  (g) For the condition of taxiing with the control locks disengaged, or where control locks are not provided, the following apply:
(effective 2017/06/19) (1) The control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads.
 (2) The parts of the control systems between the stops nearest the surfaces and the flight deck controls must be designed to the resultant limit loads, except that the parts of the control system where loads are eventually reacted by the pilot need not exceed:
 (i) the loads corresponding to the maximum pilot loads in 525.397(c) for each pilot alone; or
 (ii) 0.75 times these maximum loads for each pilot when the pilot forces are applied in the same direction.
(Change 5253 (911101))
(Change 5258)
525.427 Unsymmetrical Loads
 (a) In designing the aeroplane for lateral gust, yaw manoeuvre and roll manoeuvre conditions, account must be taken of unsymmetrical loads on the empennage arising from effects such as slipstream and aerodynamic interference with the wing, vertical fin and other aerodynamic surfaces.

(b) The horizontal tail must be assumed to be subjected to unsymmetrical loading conditions determined as follows:
 (1) 100 percent of the maximum loading from the symmetrical manoeuvre conditions of 525.331 and the vertical gust conditions of 525.341(a) acting separately on the surface on one side of the plane of symmetry; and
 (2) 80 percent of these loadings acting on the other side.
 (c) For empennage arrangements where the horizontal tail surfaces have dihedral angles greater than plus or minus 10 degrees, or are supported by the vertical tail surfaces, the surfaces and the supporting structure must be designed for gust velocities specified in 525.341(a) acting in any orientation at right angles to the flight path.
 (d) Unsymmetrical loading on the empennage arising from buffet conditions of 525.305(e) must be taken into account.
(Change 5258)
525.445 Auxiliary Aerodynamic Surfaces
 (a) When significant, the aerodynamic influence between auxiliary aerodynamic surfaces, such as outboard fins and winglets, and their supporting aerodynamic surfaces, must be taken into account for all loading conditions including pitch, roll, and yaw manoeuvres, and gusts as specified in 525.341(a) acting at any orientation at right angles to the flight path.

(b) To provide for unsymmetrical loading when outboard fins extend above and below the horizontal surface, the critical vertical surface loading (load per unit area) determined under 525.391 must also be applied as follows:
 (1) 100 percent to the area of the vertical surfaces above (or below) the horizontal surface.
 (2) 80 percent to the area below (or above) the horizontal surface.
(Change 5258)
525.457 Wing Flaps
Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical loads occurring in the conditions prescribed in 525.345, accounting for the loads occurring during transition from one flap position and airspeed to another.
525.459 Special Devices
The loading for special devices using aerodynamic surfaces (such as slats, and spoilers) must be determined from test data.
(Change 5253 (911101))
Ground Loads
525.471 General

(a) Loads and equilibrium. For limit ground loads:
 (1) Limit ground loads obtained under this subchapter are considered to be external forces applied to the aeroplane structure; and
 (2) In each specified ground load condition, the external loads must be placed in equilibrium with the linear and angular inertia loads in a rational or conservative manner.

(b) Critical centres of gravity. The critical centres of gravity within the range for which certification is requested must be selected so that the maximum design loads are obtained in each landing gear element. Fore and aft, vertical, and lateral aeroplane centres of gravity must be considered. Lateral displacements of the c.g. from the aeroplane centreline which would result in main gear loads not greater than 103 percent of the critical design load for symmetrical loading conditions may be selected without considering the effects of these lateral c.g. displacements on the loading of the main gear elements, or on the aeroplane structure provided.
 (1) The lateral displacement of the c.g. results from random passenger or cargo disposition within the fuselage or from random unsymmetrical fuel loading or fuel usage; and
 (2) Appropriate loading instructions for random disposable loads are included under the provisions of 525.1583(c)(1) to ensure that the lateral displacement of the centre of gravity is maintained within these limits.
 (c) Landing gear dimension data. Figure 1 of Appendix A contains the basic landing gear dimension data.
525.473 Landing Load Conditions and Assumptions

(a) For the landing conditions specified in 525.479 to 525.485 the aeroplane is assumed to contact the ground:
 (1) In the attitudes defined in 525.479 and 525.481;
 (2) With a limit descent velocity of 10 fps at the design landing weight (the maximum weight for landing conditions at maximum descent velocity); and
 (3) With a limit descent velocity of 6 fps at the design takeoff weight (the maximum weight for landing conditions at a reduced descent velocity).
 (4) The prescribed descent velocities may be modified if it is shown that the aeroplane has design features that make it impossible to develop these velocities.
 (b) Aeroplane lift, not exceeding aeroplane weight, may be assumed unless the presence of systems or procedures significantly affects the lift.

(c) The method of analysis of aeroplane and landing gear loads must take into account at least the following elements:
 (1) Landing gear dynamic characteristics.
 (2) Spinup and springback.
 (3) Rigid body response.
 (4) Structural dynamic response of the airframe, if significant.

(d) The landing gear dynamic characteristics shall be validated by tests as defined in paragraph 525.723(a).
(amended 2001/10/01; previous version)  (e) The coefficient of friction between the tires and the ground may be established by considering the effects of skidding velocity and tire pressure. However, this coefficient of friction need not be more than 0.8.
(Change 5258)
525.477 Landing Gear Arrangement
Sections 525.479 through 525.485 apply to aeroplanes with conventional arrangements of main and nose gears, or main and tail gears, when normal operating techniques are used.
525.479 Level Landing Conditions

(a) In the level attitude, the aeroplane is assumed to contact the ground at forward velocity components, ranging from V_{L1} to 1.25 V_{L2}, parallel to the ground under the conditions prescribed in 525.473 with:
 (1) V_{L1} equal to V_{S0} (TAS) at the appropriate landing weight and in standard sea level conditions; and
 (2) V_{L2} equal to V_{S0} (TAS) at the appropriate landing weight and altitudes in a hot day temperature of 41°F above standard.
 (3) The effects of increased contact speed must be investigated if approval of downwind landings exceeding 10 knots is requested.
 (b) For the level landing attitude for aeroplanes with tail wheels, the conditions specified in this section must be investigated with the aeroplane horizontal reference line horizontal in accordance with Figure 2 of Appendix A of this chapter.

(c) For the level landing attitude for aeroplanes with nose wheels, shown in Figure 2 of Appendix A of this chapter, the conditions specified in this section must be investigated assuming the following attitudes:
 (1) An attitude in which the main wheels are assumed to contact the ground with the nose wheel just clear of the ground; and
 (2) If reasonably attainable at the specified descent and forward velocities, an attitude in which the nose and main wheels are assumed to contact the ground simultaneously.

(d) In addition to the loading conditions prescribed in paragraph (a) of this section, but with maximum vertical ground reactions calculated from paragraph (a), the following apply:
 (1) The landing gear and directly affected attaching structure must be designed for the maximum vertical ground reaction combined with an aft acting drag component of not less than 25% of this maximum vertical ground reaction.

(2) The most severe combination of loads that are likely to arise during a lateral drift landing must be taken into account. In absence of a more rational analysis of this condition, the following must be investigated:
 (i) A vertical load equal to 75% of the maximum ground reaction of 525.473 must be considered in combination with a drag and side load of 40% and 25% respectively of that vertical load.
 (ii) The shock absorber and tire deflections must be assumed to be 75% of the deflection corresponding to the maximum ground reaction of 525.473(a)(2). This load case need not be considered in combination with flat tires.
 (3) The combination of vertical and drag components is considered to be acting at the wheel axle centreline.
(Change 5258)
525.481 Taildown Landing Conditions

(a) In the taildown attitude, the aeroplane is assumed to contact the ground at forward velocity components, ranging from V_{L1} to V_{L2}, parallel to the ground under the conditions prescribed in 525.473 with:
 (1) V_{L1} equal to V_{SO} (TAS) at the appropriate landing weight and in standard sea level conditions; and
 (2) V_{L2} equal to V_{SO} (TAS) at the appropriate landing weight and altitudes in a hotday temperature of 41° F above standard.
 (3) The combination of vertical and drag components considered to be acting at the main wheel axle centreline.

(b) For the taildown landing condition for aeroplanes with tail wheels, the main and tail wheels are assumed to contact the ground simultaneously, in accordance with Figure 3 of Appendix A. Ground reaction conditions on the tail wheel are assumed to act:
 (1) Vertically; and
 (2) Up and aft through the axle at 45 degrees to the ground line.
 (c) For the taildown landing condition for aeroplanes with nose wheels, the aeroplane is assumed to be at an attitude corresponding to either the stalling angle or the maximum angle allowing clearance with the ground by each part of the aeroplane other than the main wheels, in accordance with Figure 3 of Appendix A, whichever is less.
(Change 5258)
525.483 Onegear Landing Conditions
For the onegear landing conditions, the aeroplane is assumed to be in the level attitude and to contact the ground on one main landing gear, in accordance with Figure 4 of Appendix A of this chapter. In this attitude:
 (a) The ground reactions must be the same as those obtained on that side under 525.479(d)(1); and
 (b) Each unbalanced external load must be reacted by aeroplane inertia in a rational or conservative manner.
(Change 5258)
525.485 Side Load Conditions
In addition to 525.479(d)(2) the following conditions must be considered:
 (a) For the side load condition, the aeroplane is assumed to be in the level attitude with only the main wheels contacting the ground, in accordance with Figure 5 of Appendix A.
 (b) Side loads of 0.8 of the vertical reaction (on one side) acting inward and 0.6 of the vertical reaction (on the other side) acting outward must be combined with onehalf of the maximum vertical ground reactions obtained in the level landing conditions. These loads are assumed to be applied at the ground contact point and to be resisted by the inertia of the aeroplane. The drag loads may be assumed to be zero.
(Change 5258)
525.487 Rebound Landing Condition
 (a) The landing gear and its supporting structure must be investigated for the loads occurring during rebound of the aeroplane from the landing surface.
 (b) With the landing gear fully extended and not in contact with the ground, a load factor of 20.0 must act on the unsprung weights of the landing gear. This load factor must act in the direction of motion of the unsprung weights as they reach their limiting positions in extending with relation to the sprung parts of the landing gear.
525.489 Ground Handling Conditions
Unless otherwise prescribed, the landing gear and aeroplane structure must be investigated for the conditions in 525.491 through 525.509 with the aeroplane at the design ramp weight (the maximum weight for ground handling conditions). No wing lift may be considered. The shock absorbers and tires may be assumed to be in their static position.
525.491 Taxi, Takeoff and Landing Roll
Within the range of appropriate ground speeds and approved weights, the aeroplane structure and landing gear are assumed to be subjected to loads not less than those obtained when the aircraft is operating over the roughest ground that may reasonably be expected in normal operation.
(Change 5258)
525.493 Braked Roll Conditions
 (a) An aeroplane with a tail wheel is assumed to be in the level attitude with the load on the main wheels, in accordance with Figure 6 of Appendix A. The limit vertical load factor is 1.2 at the design landing weight, and 1.0 at the design ramp weight. A drag reaction equal to the vertical reaction multiplied by a coefficient of friction of 0.8, must be combined with the vertical ground reaction and applied at the ground contact point.

(b) For an aeroplane with a nose wheel, the limit vertical load factor is 1.2 at the design landing weight, and 1.0 at the design ramp weight. A drag reaction equal to the vertical reaction, multiplied by a coefficient of friction of 0.8, must be combined with the vertical reaction and applied at the ground contact point of each wheel with brakes. The following two attitudes, in accordance with Figure 6 of Appendix A, must be considered:
 (1) The level attitude with the wheels contacting the ground and the loads distributed between the main and nose gear. Zero pitching acceleration is assumed.
 (2) The level attitude with only the main gear contacting the ground and with the pitching moment resisted by angular acceleration.
 (c) A drag reaction lower than that prescribed in this section may be used if it is substantiated that an effective drag force of 0.8 times the vertical reaction cannot be attained under any likely loading condition.
 (d) An aeroplane equipped with a nose gear must be designed to withstand the loads arising from the dynamic pitching motion of the aeroplane due to sudden application of maximum braking force. The aeroplane is considered to be at design takeoff weight with the nose and main gears in contact with the ground, and with a steadystate vertical load factor of 1.0. The steadystate nose gear reaction must be combined with the maximum incremental nose gear vertical reaction caused by the sudden application of maximum braking force as described in paragraphs (b) and (c) of this section.
 (e) In the absence of a more rational analysis, the nose gear vertical reaction prescribed in paragraph (d) of this section must be calculated according to the following formula:
${\mathrm{V}}_{\mathrm{N}}=\frac{{\mathrm{W}}_{\mathrm{T}}}{\mathrm{A}+\mathrm{B}}\left[\mathrm{B}+\frac{\mathrm{f}\mu \mathrm{AE}}{\mathrm{A}+\mathrm{B}+\mu \mathrm{E}}\right]$
Where:
V_{N} = Nose gear vertical reaction.
W_{T} = Design takeoff weight.
A = Horizontal distance between the c.g. of the aeroplane and the nose wheel.
B = Horizontal distance between the c.g. of the aeroplane and the line joining the centres of the main wheels.
E = Vertical height of the c.g. of the aeroplane above the ground in the 1.0 g static condition.
μ = Coefficient of friction of 0.80.
f = Dynamic response factor; 2.0 is to be used unless a lower factor is substantiated. In the absence of other information, the dynamic response factor f may be defined by the equation:
$\mathrm{f}=1+\mathrm{exp}\left(\frac{\pi \xi}{\sqrt{1{\mathrm{\xi}}^{2}}}\right)$
Where:
$\mathrm{\xi}$ = is the effective critical damping ratio of the rigid body pitching mode about the main landing gear effective ground contact point.
(Change 5258)
525.495 Turning
In the static position, in accordance with Figure 7 of Appendix A, the aeroplane is assumed to execute a steady turn by nose gear steering, or by application of sufficient differential power, so that the limit load factors applied at the centre of gravity are 1.0 vertically and 0.5 laterally. The side ground reaction of each wheel must be 0.5 of the vertical reaction.
525.497 Tailwheel Yawing
 (a) A vertical ground reaction equal to the static load on the tail wheel, in combination with a side component of equal magnitude, is assumed.
 (b) If there is a swivel, the tail wheel is assumed to be swivelled 90° to the aeroplane longitudinal axis with the resultant load passing through the axle.
 (c) If there is a lock, steering device, or shimmy damper the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point.
525.499 Nosewheel Yaw and Steering
 (a) A vertical load factor of 1.0 at the aeroplane centre of gravity, and a side component at the nose wheel ground contact equal to 0.8 of the vertical ground reaction at that point are assumed.

(b) With the aeroplane assumed to be in static equilibrium with the loads resulting from the use of brakes on one side of the main landing gear, the nose gear, its attaching structure, and the fuselage structure forward of the centre of gravity must be designed for the following loads:
 (1) A vertical load factor at the centre of gravity of 1.0.
 (2) A forward acting load at the aeroplane centre of gravity of 0.8 times the vertical load on one main gear.
 (3) Side and vertical loads at the ground contact point on the nose gear that are required for static equilibrium.
 (4) A side load factor at the aeroplane centre of gravity of zero.
 (c) If the loads prescribed in paragraph (b) of this section result in a nose gear side load higher than 0.8 times the vertical nose gear load, the design nose gear side load may be limited to 0.8 times the vertical load, with unbalanced yawing moments assumed to be resisted by aeroplane inertia forces.

(d) For other than the nose gear, its attaching structure, and the forward fuselage structure the loading conditions are those prescribed in paragraph (b) of this section, except that:
 (1) A lower drag reaction may be used if an effective drag force of 0.8 times the vertical reaction cannot be reached under any likely loading condition; and
 (2) The forward acting load at the centre of gravity need not exceed the maximum drag reaction on one main gear, determined in accordance with 525.493(b).
 (e) With the aeroplane at design ramp weight, and the nose gear in any steerable position, the combined application of full normal steering torque and vertical force equal to 1.33 times the maximum static reaction on the nose gear must be considered in designing the nose gear, its attaching structure, and the forward fuselage structure.
(Change 5258)
525.503 Pivoting
 (a) The aeroplane is assumed to pivot about one side of the main gear with the brakes on that side locked. The limit vertical load factor must be 1.0 and the coefficient of friction 0.8.
 (b) The aeroplane is assumed to be in static equilibrium, with the loads being applied at the ground contact points, in accordance with Figure 8 of Appendix A.
525.507 Reversed Braking
 (a) The aeroplane must be in a three point static ground attitude. Horizontal reactions parallel to the ground and directed forward must be applied at the ground contact point of each wheel with brakes. The limit loads must be equal to 0.55 times the vertical load at each wheel or to the load developed by 1.2 times the nominal maximum static brake torque, whichever is less.
 (b) For aeroplanes with nose wheels, the pitching moment must be balanced by rotational inertia.
 (c) For aeroplanes with tail wheels, the resultant of the ground reactions must pass through the centre of gravity of the aeroplane.
525.509 Towing Loads

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

(3) With W_{T} as the design ramp weight, the towing load, F_{TOW}, is:
 (i) 0.3 W_{T} for W_{T} less than 30,000 pounds (13 600 kg);

(ii) (6 W_{T} + 450,000)/70 for W_{T} between 30,000 and 100,000 pounds (13 600 kg and 45 500 kg); and
(effective 2013/12/08)  (iii) 0.15 W_{T} for W_{T} over 100,000 pounds (13 600 kg).
 (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 follows:
Tow point  Position  Load  

Magnitude  No.  Direction  
Main gear 
0.75 F_{TOW} per main gear unit 
1  Forward, parallel to drag axis.  
2  Forward, at 30° to drag axis  
3  Aft, parallel to drag axis.  
4  Aft, at 30° to drag axis.  
Auxiliary gear  Swivelled forward.  1.0 F_{TOW} 
5 
Forward. 
6  Aft.  
Swivelled aft.  do 
7 
Forward. 

8  Aft.  
Swivelled 45° from forward.  0.5 F_{TOW} 
9 
Forward, in plane of wheel. 

10  Aft, in plane of wheel.  
Swivelled 45° from aft.  do 
11 
Forward, in plane of wheel. 

12  Aft, in plane of wheel. 
(Change 5254 (920801)
525.511 Ground Load: Unsymmetrical Loads on MultipleWheel Units

(a) General. Multiplewheel landing gear units are assumed to be subjected to the limit ground loads prescribed in this subchapter under paragraphs (b) through (f) of this section. In addition:
 (1) A tandem strut gear arrangement is a multiplewheel unit; and
 (2) In determining the total load on a gear unit with respect to the provisions of paragraphs (b) through (f) of this section, the transverse shift in the load centroid, due to unsymmetrical load distribution on the wheels, may be neglected.

(b) Distribution of limit loads to wheels; tires inflated. The distribution of the limit loads among the wheels of the landing gear must be established for each landing, taxiing, and ground handling condition, taking into account the effects of the following factors:
 (1) The number of wheels and their physical arrangements. For truck type landing gear units, the effects of any seesaw motion of the truck during the landing impact must be considered in determining the maximum design loads for the fore and aft wheel pairs.
 (2) Any differentials in tire diameters resulting from a combination of manufacturing tolerances, tire growth, and tire wear. A maximum tirediameter differential equal to 2/3 of the most unfavourable combination of diameter variations that is obtained when taking into account manufacturing tolerances, tire growth, and tire wear, may be assumed.
 (3) Any unequal tire inflation pressure, assuming the maximum variation to be +5 percent of the nominal tire inflation pressure.
 (4) A runway crown of zero and a runway crown having a convex upward shape that may be approximated by a slope of 1½ percent with the horizontal. Runway crown effects must be considered with the nose gear unit on either slope of the crown.
 (5) The aeroplane attitude.
 (6) Any structural deflections.

(c) Deflated tires. The effect of deflated tires on the structure must be considered with respect to the loading conditions specified in paragraphs (d) through (f) of this section, taking into account the physical arrangement of the gear components. In addition:
 (1) The deflation of any one tire for each multiple wheel landing gear unit, and the deflation of any two critical tires for each landing gear unit using four or more wheels per unit, must be considered; and
 (2) The ground reactions must be applied to the wheels with inflated tires except that, for multiplewheel gear units with more than one shock strut, a rational distribution of the ground reactions between the deflated and inflated tires, accounting for the differences in shock strut extensions resulting from a deflated tire, may be used.
 (d) Landing conditions. For one and for two deflated tires, the applied load to each gear unit is assumed to be 60 percent and 50 percent, respectively, of the limit load applied to each gear for each of the prescribed landing conditions. However, for the drift landing condition of 525.485, 100 percent of the vertical load must be applied.

(e) Taxiing and ground handling conditions. For one and for two deflated tires:
 (1) The applied side or drag load factor, or both factors, at the centre of gravity must be the most critical value up to 50 percent and 40 percent, respectively, of the limit side or drag load factors, or both factors, corresponding to the most severe condition resulting from the consideration of the prescribed taxiing and ground handling conditions;
 (2) For the braked roll conditions of 525.493(a) and (b)(2), the drag loads on each inflated tire may not be less than those at each tire for the symmetrical load distribution with no deflated tires;
 (3) The vertical load factor at the centre of gravity must be 60 percent and 50 percent, respectively, of the factor with no deflated tires, except that it may not be less that 1g; and
 (4) Pivoting need not be considered.
 (f) Towing conditions. For one and for two deflated tires, the towing load, F_{TOW}, must be 60 percent and 50 percent, respectively, of the load prescribed.
525.519 Jacking and Tiedown Provisions
 (a) General. The aeroplane must be designed to withstand the limit load conditions resulting from the static ground load conditions of paragraph (b) of this section and, if applicable, paragraph (c) of this section at the most critical combinations of aeroplane weight and centre of gravity. The maximum allowable load at each jack pad must be specified.

(b) Jacking. The aeroplane must have provisions for jacking and must withstand the following limit loads when the aeroplane is supported on jacks:
 (1) For jacking by the landing gear at the maximum ramp weight of the aeroplane, the aeroplane structure must be designed for a vertical load of 1.33 times the vertical static reaction at each jacking point acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.

(2) For jacking by other aeroplane structure at maximum approved jacking weight:
 (i) The aeroplane structure must be designed for a vertical load of 1.33 times the vertical reaction at each jacking point acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.
 (ii) The jacking pads and local structure must be designed for a vertical load of 2.0 times the vertical static reaction at each jacking point, acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.
 (c) Tiedown. If tiedown points are provided, the main tiedown points and local structure must withstand the limit loads resulting from a 65knot horizontal wind from any direction.
(Change 5257 (960930))
Water Loads
525.521 General
 (a) Seaplanes must be designed for the water loads developed during takeoff and landing, with the seaplane in any attitude likely to occur in normal operation, and at the appropriate forward and sinking velocities under the most severe sea conditions likely to be encountered.
 (b) Unless a more rational analysis of the water loads is made, or the standards in ANC3 are used, 525.523 through 525.537 apply.
 (c) The requirements of this section and 525.523 through 525.537 apply also to amphibians.
525.523 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 525.531, 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.
525.525 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 525.527.
 (b) In applying the loads resulting from the load factors prescribed in 525.527, 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 525.533(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 525.531, the aerodynamic lift on the seaplane during the impact is assumed to be 2/3 of the weight of the seaplane.
525.527 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
${n}_{\mathit{w}}\mathit{=}\frac{{\mathit{C}}_{\mathit{1}}{{\mathit{V}}_{\mathit{s}\mathit{o}}}^{\mathit{2}}}{{\left({\mathit{tan}}^{\mathit{\beta}}\right)}^{\mathit{2}\mathit{/}\mathit{3}}\mathit{}{\mathit{W}}^{\mathit{}\mathit{1}\mathit{/}\mathit{3}}}$
 (2) For the bow and stern landing cases
$\left({n}_{w}=\frac{{C}_{\mathit{1}}{{V}_{\mathit{s}\mathit{o}}}^{\mathit{2}}}{{\left(\mathrm{tan}\beta \right)}^{2/3}{m}^{\mathit{1}\mathit{/}\mathit{3}}}\times \frac{{K}_{\mathit{1}}}{{\left(\mathit{1}+{{r}_{\mathit{x}}}^{\mathit{2}}\right)}^{2/3}}\right)$

(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 1 of Appendix B.
 (5) W = seaplane design landing weight in pounds.
 (6) K_{1} = empirical hull station weighing factor, in accordance with Figure 2 of Appendix B.
 (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 2 of Appendix B. This reduction applies only to the design of the carry through and seaplane structure.
525.529 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 525.527. 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 to 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 at and a side load of 0.25 tan b at one float times the step landing load reached under 525.527. 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.
525.531 Hull and Main Float Takeoff Conditions
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.
$n=\frac{{C}_{TO}{{V}_{{S}_{1}}}^{2}}{\left({\mathrm{tan}}^{2/3}\beta \right){W}^{1/3}}$
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.
525.533 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 3 of Appendix B. The pressure at the keel (p.s.i.) is computed as follows:
${P}_{k}={C}_{2}\times \frac{{K}_{2}{{V}_{S1}}^{2}}{\mathrm{tan}{\beta}_{k}}$
where:
P_{K} = pressure (p.s.i.) at the keel;
C_{2} = 0.00213;
K_{2} = hull station weighing factor, in accordance with Figure 2 of Appendix B;
V_{S1} = seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and
ß_{k} = angle of dead rise at keel, in accordance with Figure 1 of Appendix B.
 (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 3 of Appendix B. The pressure distribution is the same as that prescribed in subparagraph (1) of this paragraph for an unflared bottom except that the pressure at the chine is computed as follows:
${P}_{ch}={C}_{3}\times \frac{{K}_{2}{{V}_{S1}}^{2}}{\mathrm{tan}\beta}$
where:
P_{ch} = pressure (p.s.i.) at the chine;
C_{3} = 0.0016;
K_{2} = hull station weighting factor, in accordance with Figure 2 of Appendix B;
V_{S1} = seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and
ß = angle of dead rise at appropriate station.
The area over which these pressures are applied must simulate pressures occurring during high localised 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:
$P={C}_{4}\times \frac{{K}_{2}{{V}_{S0}}^{2}}{\mathrm{tan}\beta}$
where:
P = pressure (p.s.i.);
C_{4} = 0.078 C_{1} (with C_{1} computed under 525.527);
K_{2} = hull station weighting factor, determined in accordance with Figure 2 of Appendix B;
V_{S0} = seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and
ß = angle of dead rise at appropriate station.

(2) The unsymmetrical pressure distribution consists of the pressures prescribed in subparagraph (1) of this paragraph 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 3 of Appendix B.
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.
525.535 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:
$L=\frac{{C}_{5}{{V}_{S0}}^{2}{W}^{2/3}}{\left({\mathrm{tan}}^{2/3}{\beta}_{S}\right){\left(1+{{r}_{y}}^{2}\right)}^{2/3}}$
where:
L = limit load (lbs.);
C_{5} = 0.0053;
V_{S0} = 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 loads consists of a component equal to 0.75 times the load specified in paragraph (a) of this section and a side component equal to 0.25 tan 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 b 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:
$\rho $gV
${C}_{x}\frac{\rho}{2}{V}^{2/3}{\left(K{V}_{S0}\right)}^{2}$
$\left(0,465{C}_{x}\frac{\rho}{2}{V}^{2/3}{\left(K{V}_{S0}\right)}^{2}\right)$
$\left(0.465{C}_{y}\frac{\rho}{2}{V}^{2/3}{\left(K{V}_{S0}\right)}^{2}\right)$
where:
p = 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 0.106;
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_{S0} in normal operations;
V_{S0} = 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 525.533, 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.
525.537 Seawing Loads
Seawing design loads must be based on applicable test data.
Emergency Landing Conditions
525.561 General
 (a) The aeroplane, although it may be damaged in emergency landing conditions on land or water, 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 in a minor crash landing when:
 (1) Proper use is made of seats, belts, and all other safety design provisions;
 (2) The wheels are retracted (where applicable); and

(3) The occupant experiences the following ultimate inertia forces acting separately relative to the surrounding structure:
 (i) Upward 3.0g.
 (ii) Forward 9.0g.
 (iii) Sideward, 3.0g on the airframe; and 4.0g on the seats and their attachments.
 (iv) Downward, 6.0g.
 (v) Rearward, 1.5g.

(c) For equipment, cargo in the passenger compartments and any other large masses, the following apply:

(1) Except as provided in paragraph (c)(2) of this section, these items must be positioned so that if they break loose they will be unlikely to:
 (i) Cause direct injury to occupants;
 (ii) Penetrate fuel tanks or lines or cause fire or explosion hazard by damage to adjacent systems; or
 (iii) Nullify any of the escape facilities provided for use after an emergency landing.
 (2) When such positioning is not practical (e.g. fuselage mounted engines or auxiliary power units) each such item of mass shall be restrained under all loads up to those specified in paragraph (b)(3) of this section. The local attachments for these items should be designed to withstand 1.33 times the specified loads if these items are subject to severe wear and tear through frequent removal (e.g. quick change interior items).

(1) Except as provided in paragraph (c)(2) of this section, these items must be positioned so that if they break loose they will be unlikely to:
 (d) Seats and items of mass (and their supporting structure) must not deform under any loads up to those specified in paragraph (b)(3) of this section in any manner that would impede subsequent rapid evacuation of occupants.
(Change 5252 (890101))
(Change 5258)
525.562 Emergency Landing Dynamic Conditions

(a) The seat and restraint system in the aeroplane must be designed as prescribed in this section to protect each occupant during an emergency landing condition when:
 (1) Proper use is made of seats, safety belts, and shoulder harnesses provided for in the design; and
 (2) The occupant is exposed to loads resulting from the conditions prescribed in this section.

(b) Each seat type design approved for crew or passenger occupancy during takeoff and landing must successfully complete dynamic tests or be demonstrated by rational analysis based on dynamic tests of a similar type seat, in accordance with each of the following emergency landing conditions. The tests must be conducted with an occupant simulated by a 170pound anthropomorphic test dummy, as defined by the USA 49 CFR Part 572, Subpart B, or its equivalent, sitting in the normal upright position.
 (1) A change in downward vertical velocity (D v) of not less than 35 feet per second, with the aeroplane's longitudinal axis canted downward 30 degrees with respect to the horizontal plane and with the wings level. Peak floor deceleration must occur in not more than 0.08 seconds after impact and must reach a minimum of 14g.
 (2) A change in forward longitudinal velocity (D v) of not less than 44 feet per second, with the aeroplane's longitudinal axis horizontal and yawed 10 degrees either right or left, whichever would cause the greatest likelihood of the upper torso restraint system (where installed) moving off the occupant's shoulder, and with the wings level. Peak floor deceleration must occur in not more than 0.09 seconds after impact and must reach a minimum of 16g. Where floor rails or floor fittings are used to attach the seating devices to the test fixture, the rails or fittings must be misaligned with respect to the adjacent set of rails or fittings by at least 10 degrees vertically (i.e., out of Parallel) with one rolled 10 degrees.

(c) The following performance measures must not be exceeded during the dynamic tests conducted in accordance with paragraph (b) of this section:
 (1) Where upper torso straps are used for crew members, tension loads in individual straps must not exceed 1,750 pounds. If dual straps are used for restraining the upper torso, the total strap tension loads must not exceed 2,000 pounds.
 (2) The maximum compressive load measured between the pelvis and the lumbar column of the anthropomorphic dummy must not exceed 1,500 pounds.
 (3) The upper torso restraint straps (where installed) must remain on the occupant's shoulder during the impact.
 (4) The lap safety belt must remain on the occupant's pelvis during the impact.
 (5) Each occupant must be protected from serious head injury under the conditions prescribed in paragraph (b) of this section. Where head contact with seats or other structure can occur, protection must be provided so that the head impact does not exceed a Head Injury Criterion (HIC) of 1,000 units. The level of HIC is defined by the equation:
$\mathrm{HIC}={\left\{\left({\mathrm{t}}_{2}{\mathrm{t}}_{1}\right){\left[\frac{1}{\left({\mathrm{t}}_{2}{\mathrm{t}}_{1}\right)}{\int}_{{\mathrm{t}}_{1}}^{{\mathrm{t}}_{2}}\mathrm{a}\left(\mathrm{t}\right)\mathrm{dt}\right]}^{2.5}\right\}}_{\mathrm{max}}$
Where:
t_{1} is the initial integration time,
t_{2} is the final integration time, and
a(t) is the total acceleration vs. time curve for the head strike, and where:
(t) is in seconds, and (a) is in units of gravity (g).
 (6) Where leg injuries may result from contact with seats or other structure, protection must be provided to prevent axially compressive loads exceeding 2,250 pounds in each femur.
 (7) The seat must remain attached at all points of attachment, although the structure may have yielded.
 (8) Seats must not yield under the tests specified in paragraphs (b)(1) and (b)(2) of this section to the extent they would impede rapid evacuation of the aeroplane occupants.
(Change 5252 (890101))
525.563 Structural Ditching Provisions
Structural strength considerations of ditching provisions must be in accordance with 525.801(e).
Fatigue Evaluation
525.571 Damagetolerance and Fatigue Evaluation of Structure

(a) General. An evaluation of the strength, detail design, and fabrication must show that catastrophic failure due to fatigue, corrosion, manufacturing defects, or accidental damage, will be avoided throughout the operational life of the aeroplane. This evaluation must be conducted in accordance with the provisions of paragraphs (b) and (e) of this section, except as specified in paragraph (c) of this section, for each part of the structure that could contribute to a catastrophic failure (such as wing, empennage, control surfaces and their systems, the fuselage, engine mounting, landing gear, and their related primary attachments). For turbojet powered aeroplanes, those parts that could contribute to a catastrophic failure must also be evaluated under paragraph (d) of this section. In addition, the following apply:

(1) Each evaluation required by this section must include:
 (i) The typical loading spectra, temperatures, and humidities expected in services;
 (ii) The identification of principal structural elements and detail design points, the failure of which could cause catastrophic failure of the aeroplane; and
 (iii) An analysis, supported by test evidence, of the principal structural elements and detail design points identified in paragraph (a)(1)(ii) of this section.
 (2) The service history of aeroplanes of similar structural design, taking due account of differences in operating conditions and procedures, may be used in the evaluations required by this section.

(3) Based on the evaluations required by this section, inspections or other procedures must be established, as necessary, to prevent catastrophic failure, and must be included in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by 525.1529. The limit of validity of the engineering data that supports the structural maintenance program (hereafter referred to as LOV), stated as a number of total accumulated flight cycles or flight hours or both, established by this section must also be included in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by 525.1529. Inspection thresholds for the following types of structure must be established based on crack growth analyses and/or test, assuming the structure contains an initial flaw of the maximum probable size that could exist as a result of manufacturing or service induced damage:
(amended 2012/03/27; previous version) (i) Single load path structure, and
 (ii) Multiple load path "failsafe" structure and crack arrest "failsafe" structure, where it cannot be demonstrated that load path failure, partial failure, or crack arrest will be detected and repaired during normal ,maintenance, inspection, or operation of an aeroplane prior to failure of the remaining structure.

(1) Each evaluation required by this section must include:

(b) Damagetolerance evaluation. The evaluation must include a determination of the probable locations and modes of damage due to fatigue, corrosion, or accidental damage. Repeated load and static analyses supported by test evidence and (if available) service experience must also be incorporated in the evaluation. Special consideration for widespread fatigue damage must be included where the design is such that this type of damage could occur. An LOV must be established that corresponds to the period of time, stated as a number of total accumulated flight cycles or flight hours or both, during which it is demonstrated that widespread fatigue damage will not occur in the aeroplane structure. This demonstration must be by fullscale fatigue test evidence. The type certificate may be issued prior to completion of the full‑scale fatigue testing, provided the Minister has approved a plan for completing the required tests, and the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by 525.1529 of this chapter specifies that no aeroplane may be operated beyond a number of cycles equal to ½ the number of cycles accumulated on the fatigue test article, until such testing is completed. 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 loads (considered as static ultimate loads) corresponding to the following conditions:
(amended 2012/03/27; previous version) (1) The limit symmetrical manoeuvring conditions specified in 525.337 at all speeds up to V_{C} and in 525.345.
 (2) The limit gust conditions specified in 525.341 at the specified speeds up to V_{C}, and in 525.345.
 (3) The limit rolling conditions specified in 525.349 and the limit unsymmetrical conditions specified in 525.367 and 525.427(a) through (c), at speeds up to V_{C}.
 (4) The limit yaw manoeuvring conditions specified in 525.351(a) at the specified speeds up to V_{C}.

(5) For pressurised cabins, the following conditions:
 (i) The normal operating differential pressure combined with the expected external aerodynamic pressures applied simultaneously with the flight loading conditions specified in paragraphs (b)(1) through (4) of this section, if they have a significant effect.
 (ii) The maximum value of normal operating differential pressure (including the expected external aerodynamic pressures during 1g level flight) multiplied by a factor of 1.15, omitting other loads.

(6) For landing gear and directly affected airframe structure, the limit ground loading conditions specified in 525.473, 525.491, and 525.493.
If significant changes in structural stiffness or geometry, or both, follow from a structural failure, or partial failure, the effect on damage tolerance must be further investigated.
 (c) Fatigue (safelife) evaluation. Compliance with the damagetolerance requirements of paragraph (b) of this section is not required if the applicant establishes that their application for particular structure is impractical. 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.
 (d) Sonic fatigue strength. It must be shown by analysis, supported by test evidence, or by the service history of aeroplanes of similar structural design and sonic excitation environment, that:
 (1) Sonic fatigue cracks are not probable in any part of the flight structure subject to sonic excitation; or
 (2) Catastrophic failure caused by sonic cracks is not probable assuming that the loads prescribed in paragraph (b) of this section are applied to all areas affected by those cracks.

(e) Damagetolerance (discrete source) evaluation. The aeroplane must be capable of successfully completing a flight during which likely structural damage occurs as a result of:
 (1) Impact with a 4pound bird when the velocity of the aeroplane relative to the bird along the aeroplane's flight path is equal to V_{C} at sea level or 0.85 V_{C} at 8,000 feet, whichever is more critical;
 (2) Uncontained fan blade impact;
 (3) Uncontained engine failure; or

(4) Uncontained high energy rotating machinery failure.
The damaged structure must be able to withstand the static loads (considered as ultimate loads) which are reasonably expected to occur on the flight. Dynamic effects on these static loads need not be considered. Corrective action to be taken by the pilot following the incident, such as limiting manoeuvres, avoiding turbulence, and reducing speed, must be considered. If significant changes in structural stiffness or geometry, or both, follow from a structural failure or partial failure, the effect on damage tolerance must be further investigated.
(Change 5251 (870101))
(Change 5252 (890101))
(Change 5253 (911101))
(Change 5258)
525.573 (Reserved)
Lightning Protection
525.581 Lightning Protection
 (a) The aeroplane must be protected against catastrophic effects from lightning.

(b) For metallic components, compliance with paragraph (a) of this section may be shown by:
 (1) Bonding the components properly to the airframe; or
 (2) Designing the components so that a strike will not endanger the aeroplane.

(c) For nonmetallic components, compliance with paragraph (a) of this section may be shown by:
 (1) Designing the components to minimise the effect of a strike; or
 (2) Incorporating acceptable means of diverting the resulting electrical current so as not to endanger the aeroplane.
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