Certification Specifications for Large Aeroplanes（CS-25）

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Certification Specifications for Large Aeroplanes（CS-25）

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1

498w

a

c

V

ref

U

g

K

1 V s1

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

+

where –

Vsl = the 1-g stalling speed based on CNAmax

with the flaps retracted at the particular weight under

consideration;

CNAmax = the maximum aeroplane normal force

coefficient;

Vc = design cruise speed (knots equivalent

airspeed);

Uref = the reference gust velocity (feet per

second equivalent airspeed) from CS 25.341(a)(5)(i);

w = average wing loading (pounds per

square foot) at the particular weight

under consideration.

Kg =

µ 5.3

.88µ

+

µ=

cag ρ

w 2

ρ= 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, CNA per radian;

(2) At altitudes where Vc is limited by

Mach number –

(i) VB may be chosen to provide an

optimum margin between low and high

speed buffet boundaries; and,

(ii) VB need not be greater than VC.

(e) Design wing-flap speeds, VF. For VF, the

following apply:

(1) The design wing-flap speed for each

wing-flap position (established in accordance with

CS 25.697 (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

wing-flap position to another.

(2) If an automatic wing-flap positioning or

load limiting device is used, the speeds and

corresponding wing-flap positions programmed or

allowed by the device may be used.

(3) VF may not be less than –

(i) 1·6 VS1 with the wing-flaps in

take-off position at maximum take-off

weight;

(ii) 1·8 VS1 with the wing-flaps in

approach position at maximum landing

weight; and

(iii) 1·8 VS0 with the wing-flaps in

landing position at maximum landing

weight.

(f) Design drag device speeds, VDD. 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, VDD may not

be less than VD. 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.

1-C-5

Annex to ED Decision 2008/006/R

Amendment 5

CS–25 BOOK 1

CS 25.337 Limit manoeuvring load

factors

(See AMC 25.337)

(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

paragraph. Pitching velocities appropriate to the

corresponding pull-up and steady turn manoeuvres

must be taken into account.

(b) The positive limit manoeuvring load factor

‘n’ for any speed up to VD may not be less than 2·1 +

24 000

W +10 000

⎛

⎝

⎜

⎞

⎠

⎟except that ‘n’ may not be less than

2·5 and need not be greater than 3·8 – where ‘W’ is

the design maximum take-off weight (lb).

(c) The negative limit manoeuvring load factor

–

(1) May not be less than –1·0 at speeds up

to VC; and

(2) Must vary linearly with speed from

the value at VC to zero at VD.

(d) Manoeuvring load factors lower than those

specified in this paragraph may be used if the

aeroplane has design features that make it impossible

to exceed these values in flight.

CS 25.341 Gust and turbulence loads

(See AMC 25.341)

(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

following 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 taken as

follows:

⎥

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(iii) In addition, for cases where the

aeroplane response to the specified

cockpit pitch control motion does not

achieve the prescribed limit load factors

then the following cockpit pitch control

motion must be used:

δ(t) = δ1 sin(ωt) for 0 ≤t ≤t1

δ(t) = δ1 for t1 ≤t ≤t2

δ(t) = δ1 sin(ω[t + t1 - t2]) for t2 ≤t ≤tmax

where:

t1 = π/2ω

t2 = t1 + Δt

tmax = t2 + 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).

time

δ

Cockpit Control

deflection

t t 2 1

Δt

1

δ

tmax

δ

1

−

(iv) In cases where the cockpit 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 must be taken into account.

(v) Aeroplane loads that occur

beyond the following times need not be

considered:

(A) For the nose-up pitching

manoeuvre, the time at which the

normal acceleration at the c.g. goes

below 0g;

(B) For the nose-down

pitching manoeuvre, the time at

which the normal acceleration at the

c.g. goes above the positive limit

load factor prescribed in CS 25.337;

(C) tmax.

1-C-3

Annex to ED Decision 2008/006/R

Amendment 5

CS–25 BOOK 1

CS 25.333 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 subparagraph (b) of this paragraph. This envelope must

also be used in determining the aeroplane structural

operating limitations as specified in CS 25.1501.

(b) Manoeuvring envelope

CS 25.335 Design airspeeds

The selected design airspeeds are equivalent

airspeeds (EAS). Estimated values of VS0 and VS1

must be conservative.

(a) Design cruising speed, VC. For VC, the

following apply:

(1) The minimum value of VC must be

sufficiently greater than VB to provide for

inadvertent speed increases likely to occur as a

result of severe atmospheric turbulence.

(2) Except as provided in sub-paragraph

25.335(d)(2), VC may not be less than VB + 1·32

Uref (with Uref as specified in sub-paragraph

25.341(a)(5)(i). However, VC need not exceed the

maximum speed in level flight at maximum

continuous power for the corresponding altitude.

(3) At altitudes where VD is limited by

Mach number, VC may be limited to a selected

Mach number. (See CS 25.1505.)

(b) Design dive speed, VD. VD must be

selected so that VC/MC is not greater than 0·8 VD/MD,

or so that the minimum speed margin between VC/MC

and VD/MD is the greater of the following values:

(1) From an initial condition of stabilised

flight at VC/MC, the aeroplane is upset, flown for

20 seconds along a flight path 7·5º below the

initial path, and then pulled up at a load factor of

1·5 g (0·5 g acceleration increment). The speed

increase occurring in this manoeuvre may be

calculated if reliable or conservative aerodynamic

data issued. Power as specified in CS 25.175

(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;

1-C-4

Annex to ED Decision 2008/006/R

Amendment 5

CS–25 BOOK 1

(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 MC 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. (See AMC

25.335(b)(2))

(c) Design manoeuvring speed, VA. For VA,

the following apply:

(1) VA may not be less than VS1 n where

–

(i) n is the limit positive

manoeuvring load factor at VC; and

(ii) VS1 is the stalling speed with

wing-flaps retracted.

(2) VA and VS must be evaluated at the

design weight and altitude under consideration.

(3) VA need not be more than VC or the

speed at which the positive CNmax curve intersects

the positive manoeuvre load factor line, whichever

is less.

(d) Design speed for maximum gust intensity,

VB.

(1) VB may not be less than

2

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(2) In determining elevator angles and

chordwise load distribution in the manoeuvring

conditions of sub-paragraphs (b) and (c) of this

paragraph, the effect of corresponding pitching

velocities must be taken into account. The in-trim

and out-of-trim flight conditions specified in CS

25.255 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

CS 25.333 (b) must be investigated.

(c) Manoeuvring pitching conditions. The

following conditions must be investigated:

(1) Maximum pitch control displacement

at VA. The aeroplane is assumed to be flying in

steady level flight (point A1, CS 25.333 (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 which occur subsequent to the time when

normal acceleration at the c.g. exceeds the

positive limit manoeuvring load factor (at point

A2 in CS.333(b)), or the resulting tailplane normal

load reaches its maximum, whichever occurs first,

need not be considered.

(2) Checked manoeuvre between VA and

VD. Nose up checked pitching manoeuvres must

be analysed in which the positive limit load

factor prescribed in CS 25.337 is achieved. As

a separate condition, nose down checked

pitching manoeuvres must be analysed in which

a limit load factor of 0 is achieved. In defining

the aeroplane loads the cockpit pitch control

motions described in sub-paragraphs (i), (ii),

(iii) and (iv) of this paragraph must be used:

(i) The aeroplane is assumed to be

flying in steady level flight at any speed

between VA and VD and the cockpit pitch

control is moved in accordance with the

following formula:

1-C-2

Annex to ED Decision 2008/006/R

Amendment 5

CS–25 BOOK 1

δ(t) = δ1 sin(ωt) for 0 t tmax ≤≤

where:

δ1 = the maximum available

displacement of the cockpit pitch

control in the initial direction, as

limited by the control system

stops, control surface stops, or by

pilot effort in accordance with

CS 25.397(b);

δ(t) = the displacement of the cockpit

pitch control as a function of

time. In the initial direction δ(t)

is limited to δ1. In the reverse

direction, δ(t) may be truncated

at the maximum available

displacement of the cockpit pitch

control as limited by the control

system stops, control surface

stops, or by pilot effort in

accordance with CS 25.397(b);

tmax = 3π/2ω;

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:

ω

π

=

V

VA 2 radians per second;

where:

V = the speed of the aeroplane at

entry to the manoeuvre.

VA = the design manoeuvring

speed prescribed in

CS 25.335(c)

(ii) For nose-up pitching

manoeuvres the complete cockpit pitch

control displacement history may be

scaled down in amplitude to the extent

just necessary to ensure that the positive

limit load factor prescribed in CS 25.337

is not exceeded. For nose-down pitching

manoeuvres the complete cockpit control

displacement history may be scaled down

in amplitude to the extent just necessary

to ensure that the normal acceleration at

the c.g. does not go below 0g.

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CS 25.305 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

3-second 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) Reserved

(e) The aeroplane must be designed to

withstand any vibration and buffeting that might

occur in any likely operating condition up to VD/MD,

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 Agency.

(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 loads must be treated in accordance

with the requirements of CS 25.302.

[Amdt. No.:25/1]

CS 25.307 Proof of structure

(See AMC 25.307)

(a) Compliance with the strength and

deformation requirements of this Subpart must be

shown for each critical loading condition. Structural

analysis may be used only if the structure conforms

to that for which experience has shown this method

to be reliable. In other cases, substantiating tests

must be made to load levels that are sufficient to

verify structural behaviour up to loads specified in

CS 25.305.

(b) Reserved

(c) Reserved

1-C-1

Annex to ED Decision 2008/006/R

Amendment 5

SUBPART C – STRUCTURE

CS–25 BOOK 1

(d) When static or dynamic tests are used to show

compliance with the requirements of CS 25.305 (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.

[Amdt. No.:25/1]

FLIGHT LOADS

CS 25.321 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 Subpart 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.

FLIGHT MANOEUVRE AND GUST CONDITIONS

CS 25.331 Symmetric manoeuvring

conditions

(a) Procedure. For the analysis of the

manoeuvring flight conditions specified in subparagraphs (b) and (c) of this paragraph, 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.