Certification Specifications for Large Aeroplanes（CS-25）

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(c) Longitudinal trim. The aeroplane must

maintain longitudinal trim during –

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(1) A climb with maximum continuous

power at a speed not more than 1·3 VSR1 , with the

landing gear retracted, and the wing-flaps (i)

retracted and (ii) in the take-off position;

(2) Either a glide with power off at a

speed not more than 1·3 VSR1 , or an approach

within the normal range of approach speeds

appropriate to the weight and configuration with

power settings corresponding to a 3º glidepath,

whichever is the most severe, with the landing

gear extended, the wing-flaps retracted and

extended, and with the most unfavourable

combination of centre of gravity position and

weight approved for landing; and

(3) Level flight at any speed from

1·3 VSR1 , to VMO/MMO, with the landing gear and

wing-flaps retracted, and from 1·3 VSR1 to VLE

with the landing gear extended.

(d) Longitudinal, directional, and lateral trim.

The aeroplane must maintain longitudinal,

directional, and lateral trim (and for lateral trim, the

angle of bank may not exceed 5º) at 1·3 VSR1 , during

the climbing flight with –

(1) The critical engine inoperative;

(2) The remaining engines at maximum

continuous power; and

(3) The landing gear and wing-flaps

retracted.

(e) Aeroplanes with four or more engines. Each

aeroplane with four or more engines must also

maintain trim in rectilinear flight with the most

unfavourable centre of gravity and at the climb

speed, configuration, and power required by CS

25.123 (a) for the purpose of establishing the enroute flight path with two engines inoperative.

STABILITY

CS 25.171 General

The aeroplane must be longitudinally, directionally

and laterally stable in accordance with the provisions

of CS 25.173 to 25.177. In addition, suitable stability

and control feel (static stability) is required in any

condition normally encountered in service, if flight

tests show it is necessary for safe operation.

CS 25.173 Static longitudinal stability

Under the conditions specified in CS 25.175, the

characteristics of the elevator control forces

(including friction) must be as follows:

(a) A pull must be required to obtain and

maintain speeds below the specified trim speed, and a

push must be required to obtain and maintain speeds

above the specified trim speed. This must be shown

at any speed that can be obtained except speeds

higher than the landing gear or wing flap operating

limit speeds or VFC/MFC, whichever is appropriate, or

lower than the minimum speed for steady unstalled

flight.

(b) The airspeed must return to within 10% of

the original trim speed for the climb, approach and

landing conditions specified in CS 25.175 (a), (c) and

(d), and must return to within 7·5% of the original

trim speed for the cruising condition specified in CS

25.175 (b), when the control force is slowly released

from any speed within the range specified in subparagraph (a) of this paragraph.

(c) The average gradient of the stable slope of

the stick force versus speed curve may not be less

than 4 N (1 pound) for each 11,2 km/h (6 kt). (See

AMC 25.173(c).)

(d) Within the free return speed range specified

in sub-paragraph (b) of this paragraph, it is

permissible for the aeroplane, without control forces,

to stabilise on speeds above or below the desired trim

speeds if exceptional attention on the part of the pilot

is not required to return to and maintain the desired

trim speed and altitude.

CS 25.175 Demonstration of static

longitudinal stability

Static longitudinal stability must be shown as

follows:

(a) Climb. The stick force curve must have a

stable slope at speeds between 85% and 115% of the

speed at which the aeroplane –

(1) Is trimmed with –

(i) Wing-flaps retracted;

(ii) Landing gear retracted;

(iii) Maximum take-off weight; and

(iv) The maximum power or thrust

selected by the applicant as an operating

limitation for use during climb; and

(2) Is trimmed at the speed for best rateof-climb except that the speed need not be less

than 1·3 VSR1 .

(b) Cruise. Static longitudinal stability must be

shown in the cruise condition as follows:

(1) With the landing gear retracted at high

speed, the stick force curve must have a stable

slope at all speeds within a range which is the

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greater of 15% of the trim speed plus the resulting

free return speed range, or 93 km/h (50 kt) plus

the resulting free return speed range, above and

below the trim speed (except that the speed range

need not include speeds less than 1·3 VSR1 nor

speeds greater than VFC/MFC, nor speeds that

require a stick force of more than 222 N (50 lbf)),

with –

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(i) The wing-flaps retracted;

(ii) The centre of gravity in the

most adverse position (see CS 25.27);

(iii) The most critical weight

between the maximum take-off and

maximum landing weights;

(iv) The maximum cruising power

selected by the applicant as an operating

limitation (see CS 25.1521), except that the

power need not exceed that required at

VMO/MMO; and

(v) The aeroplane trimmed for level

flight with the power required in subparagraph (iv) above.

(2) With the landing gear retracted at low

speed, the stick force curve must have a stable

slope at all speeds within a range which is the

greater of 15% of the trim speed plus the resulting

free return speed range, or 93 km/h (50 kt) plus

the resulting free return speed range, above and

below the trim speed (except that the speed range

need not include speeds less than 1·3 VSR1 nor

speeds greater than the minimum speed of the

applicable speed range prescribed in subparagraph (b)(1) of this paragraph, nor speeds that

require a stick force of more than 222 N (50 lbf)),

with –

(i) Wing-flaps, centre of gravity

position, and weight as specified in subparagraph (1) of this paragraph;

(ii) Power required for level flight

at a speed equal to

2

SR1

1·3V VMO +

; and

(iii) The aeroplane trimmed for level

flight with the power required in subparagraph (ii) above.

(3) With the landing gear extended, the

stick force curve must have a stable slope at all

speeds within a range which is the greater of 15%

of the trim speed plus the resulting free return

speed range or 93 km/h (50 kt) plus the resulting

free return speed range, above and below the trim

speed (except that the speed range need not

include speeds less than 1·3 VSR1 , nor speeds

greater than VLE, nor speeds that require a stick

force of more than 222 N (50 lbf)), with –

(i) Wing-flap, centre of gravity

position, and weight as specified in subparagraph (b)(1) of this paragraph;

(ii) The maximum cruising power

selected by the applicant as an operating

limitation, except that the power need not

exceed that required for level flight at VLE;

and

(iii) The aeroplane trimmed for level

flight with the power required in subparagraph (ii) above.

(c) Approach. The stick force curve must have

a stable slope at speeds between VSW, and 1·7 VSR1

with –

(1) Wing-flaps in the approach position;

(2) Landing gear retracted;

(3) Maximum landing weight; and

(4) The aeroplane trimmed at 1·3 VSR1 ,

with enough power to maintain level flight at this

speed.

(d) Landing. The stick force curve must have a

stable slope and the stick force may not exceed 356

N (80 lbf) at speeds between VSW, and 1·7 VSR0 with –

(1) Wing-flaps in the landing position;

(2) Landing gear extended;

(3) Maximum landing weight;

(4) The aeroplane trimmed at 1·3 VSR0

with –

(i) Power or thrust off, and

(ii) Power or thrust for level flight.

CS 25.177 Static directional and

lateral stability

(a) The static directional stability (as shown by

the tendency to recover from a skid with the rudder

free) must be positive for any landing gear and flap

position and symmetrical power condition, at speeds

from 1·13 VSR1 , up to VFE, VLE, or VFC/MFC (as

appropriate).

(b) The static lateral stability (as shown by the

tendency to raise the low wing in a sideslip with the

aileron controls free) for any landing gear and wingflap position and symmetric power condition, may

not be negative at any airspeed (except that speeds

higher than VFE need not be considered for wingflaps extended configurations nor speeds higher than

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VLE for landing gear extended configurations) in the

following airspeed ranges (see AMC 25.177(b)):

(1) From 1·13 VSR1 to VMO/MMO..

(2) From VMO/MMO to VFC/MFC, unless

the divergence is –

(i) Gradual;

(ii) Easily recognisable by the pilot;

and

(iii) Easily controllable by the pilot

(c) In straight, steady, sideslips over the range

of sideslip angles appropriate to the operation of the

aeroplane, but not less than those obtained with onehalf of the available rudder control input or a rudder

control force of 801 N (180 lbf) , the aileron and

rudder control movements and forces must be

substantially proportional to the angle of sideslip in a

stable sense; and the factor of proportionality must

lie between limits found necessary for safe operation

This requirement must be met for the configurations

and speeds specified in sub-paragraph (a) of this

paragraph. (See AMC 25.177(c).)

(d) For sideslip angles greater than those

prescribed by sub-paragraph (c) of this paragraph, up

to the angle at which full rudder control is used or a

rudder control force of 801 N (180 lbf) is obtained,

the rudder control forces may not reverse, and

increased rudder deflection must be needed for

increased angles of sideslip. Compliance with this

requirement must be shown using straight, steady

sideslips, unless full lateral control input is achieved

before reaching either full rudder control input or a

rudder control force of 801 N (180 lbf) ; a straight,

steady sideslip need not be maintained after

achieving full lateral control input. This requirement

must be met at all approved landing gear and wingflap positions for the range of operating speeds and

power conditions appropriate to each landing gear

and wing-flap position with all engines operating.

(See AMC 25.177(d).)

CS 25.181 Dynamic stability

(See AMC 25.181)

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(a) Any short period oscillation, not including

combined lateral-directional oscillations, occurring

between 1·13 VSR and maximum allowable speed

appropriate to the configuration of the aeroplane

must be heavily damped with the primary controls –

(1) Free; and

(2) In a fixed position.

(b) Any combined lateral-directional

oscillations (‘Dutch roll’) occurring between 1·13 VSR

and maximum allowable speed appropriate to the

configuration of the aeroplane must be positively

damped with controls free, and must be controllable

with normal use of the primary controls without

requiring exceptional pilot skill.

STALLS

CS 25.201 Stall demonstration

(a) Stalls must be shown in straight flight and

in 30º banked turns with –

(1) Power off; and

(2) The power necessary to maintain level

flight at 1·5 VSR1 (where VSR1 corresponds to the

reference stall speed at maximum landing weight

with flaps in the approach position and the

landing gear retracted. (See AMC 25.201(a)(2).)

(b) In each condition required by sub-paragraph

(a) of this paragraph, it must be possible to meet the

applicable requirements of CS25.203 with –

(1) Flaps, landing gear and deceleration

devices in any likely combination of positions

approved for operation; (See AMC 25.201(b)(1).)

(2) Representative weights within the

range for which certification is requested;

(3) The most adverse centre of gravity for

recovery; and

(4) The aeroplane trimmed for straight

flight at the speed prescribed in CS 25.103 (b)(6).

(c) The following procedures must be used to

show compliance with CS 25.203 :

(1) Starting at a speed sufficiently above

the stalling speed to ensure that a steady rate of

speed reduction can be established, apply the

longitudinal control so that the speed reduction

does not exceed 0.5 m/s

2

(one knot per second)

until the aeroplane is stalled. (See AMC

25.103(c).)

(2) In addition, for turning flight stalls,

apply the longitudinal control to achieve airspeed

deceleration rates up to 5,6 km/h (3 kt) per

second. (See AMC 25.201(c)(2).)

(3) As soon as the aeroplane is stalled,

recover by normal recovery techniques.

(d) The aeroplane is considered stalled when

the behaviour of the aeroplane gives the pilot a clear

and distinctive indication of an acceptable nature that

the aeroplane is stalled. (See AMC 25.201 (d).)

Acceptable indications of a stall, occurring either

individually or in combination, are –

(1) A nose-down pitch that cannot be

readily arrested;

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(2) Buffeting, of a magnitude and severity

that is a strong and effective deterrent to further

speed reduction; or

(3) The pitch control reaches the aft stop

and no further increase in pitch attitude occurs

when the control is held full aft for a short time

before recovery is initiated. (See AMC

25.201(d)(3).)

CS 25.203 Stall characteristics

(See AMC 25.203.)

(a) It must be possible to produce and to correct

roll and yaw by unreversed use of aileron and rudder

controls, up to the time the aeroplane is stalled. No

abnormal nose-up pitching may occur. The

longitudinal control force must be positive up to and

throughout the stall. In addition, it must be possible

to promptly prevent stalling and to recover from a

stall by normal use of the controls.

(b) For level wing stalls, the roll occurring

between the stall and the completion of the recovery

may not exceed approximately 20º.

(c) For turning flight stalls, the action of the

aeroplane after the stall may not be so violent or

extreme as to make it difficult, with normal piloting

skill, to effect a prompt recovery and to regain

control of the aeroplane. The maximum bank angle

that occurs during the recovery may not exceed –

(1) Approximately 60º in the original

direction of the turn, or 30º in the opposite

direction, for deceleration rates up to 0.5 m/s

2

(1

knot per second); and

(2) Approximately 90º in the original

direction of the turn, or 60º in the opposite

direction, for deceleration rates in excess of 0.5

m/s

2

(1 knot per second).

CS 25.207 Stall warning

(a) Stall warning with sufficient margin to

prevent inadvertent stalling with the flaps and

landing gear in any normal position must be clear and

distinctive to the pilot in straight and turning flight.

(b) The warning must be furnished either

through the inherent aerodynamic qualities of the

aeroplane or by a device that will give clearly

distinguishable indications under expected conditions

of flight. However, a visual stall warning device that

requires the attention of the crew within the cockpit

is not acceptable by itself. If a warning device is

used, it must provide a warning in each of the

aeroplane configurations prescribed in sub-paragraph

(a) of this paragraph at the speed prescribed in subparagraphs (c) and (d) of this paragraph. Except for

the stall warning prescribed in paragraph (h)(2)(ii) of

this section, the stall warning for flight in icing

conditions prescribed in paragraph (e) of this section

must be provided by the same means as the stall

warning for flight in non-icing conditions. (See AMC

25.207(b).)

(c) When the speed is reduced at rates not

exceeding 0.5 m/s

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2

(one knot per second), stall

warning must begin, in each normal configuration, at

a speed, VSW, exceeding the speed at which the stall

is identified in accordance with CS 25.201 (d) by not

less than 9.3 km/h (five knots) or five percent CAS,

whichever is greater. Once initiated, stall warning

must continue until the angle of attack is reduced to

approximately that at which stall warning began.

(See AMC 25.207(c) and (d)).

(d) In addition to the requirement of subparagraph(c) of this paragraph, when the speed is

reduced at rates not exceeding 0.5 m/s

2

(one knot per

second), in straight flight with engines idling and at

the centre-of-gravity position specified in CS

25.103(b)(5), VSW, in each normal configuration,

must exceed VSR by not less than 5.6 km/h (three

knots) or three percent CAS, whichever is greater.

(See AMC 25.207(c) and (d)).

(e) In icing conditions, the stall warning margin

in straight and turning flight must be sufficient to

allow the pilot to prevent stalling (as defined in CS

25.201(d)) when the pilot starts a recovery

manoeuvre not less than three seconds after the onset

of stall warning. When demonstrating compliance

with this paragraph, the pilot must perform the

recovery manoeuvre in the same way as for the

airplane in non-icing conditions. Compliance with

this requirement must be demonstrated in flight with

the speed reduced at rates not exceeding 0.5 m/sec

2

(one knot per second), with –

(1) The more critical of the takeoff ice and

final takeoff ice accretions defined in appendix C

for each configuration used in the takeoff phase

of flight;

(2) The en route ice accretion defined in

appendix C for the en route configuration;

(3) The holding ice accretion defined in

appendix C for the holding configuration(s);

(4) The approach ice accretion defined in

appendix C for the approach configuration(s); and

(5) The landing ice accretion defined in

appendix C for the landing and go-around

configuration(s).

(f) The stall warning margin must be sufficient

in both non-icing and icing conditions to allow the

pilot to prevent stalling when the pilot starts a

recovery manoeuvre not less than one second after

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the onset of stall warning in slow-down turns with at

least 1.5g load factor normal to the flight path and

airspeed deceleration rates of at least 1 m/s

2

(2 knots

per second). When demonstrating compliance with

this paragraph for icing conditions, the pilot must

perform the recovery manoeuvre in the same way as

for the airplane in non-icing conditions. Compliance

with this requirement must be demonstrated in flight

with –

(1) The flaps and landing gear in any

normal position;

(2) The aeroplane trimmed for straight

flight at a speed of 1.3 VSR; and

(3) The power or thrust necessary to

maintain level flight at 1.3 VSR.

(g) Stall warning must also be provided in each

abnormal configuration of the high lift devices that is

likely to be used in flight following system failures

(including all configurations covered by Aeroplane

Flight Manual procedures).

(h) For flight in icing conditions before the ice

protection system has been activated and is

performing its intended function, the following

requirements apply, with the ice accretion defined in

appendix C, part II(e):

(1) If activating the ice protection system

depends on the pilot seeing a specified ice

accretion on a reference surface (not just the first

indication of icing), the requirements of this

section apply, except for paragraphs (c) and (d).

(2) For other means of activating the ice

protection system, the stall warning margin in

straight and turning flight must be sufficient to

allow the pilot to prevent stalling without

encountering any adverse flight characteristics

when the speed is reduced at rates not exceeding

0.5 m/sec

2

(one knot per second) and the pilot

performs the recovery manoeuvre in the same

way as for flight in non-icing conditions.

(i) If stall warning is provided by

the same means as for flight in non-icing

conditions, the pilot may not start the

recovery manoeuvre earlier than one second

after the onset of stall warning.

(ii) If stall warning is provided by a

different means than for flight in non-icing

conditions, the pilot may not start the

recovery manoeuvre earlier than 3 seconds

after the onset of stall warning. Also,

compliance must be shown with CS 25.203

using the demonstration prescribed by CS

25.201, except that the deceleration rates of

CS 25.201(c)(2) need not be demonstrated.

[Amdt. No.:25/3]

GROUND HANDLING CHARACTERISTICS

CS 25.231 Longitudinal stability and

control

(a) Aeroplanes may have no uncontrollable

tendency to nose over in any reasonably expected

operating condition or when rebound occurs during

landing or take-off. In addition –

(1) Wheel brakes must operate smoothly

and may not cause any undue tendency to nose

over; and

(2) If a tail-wheel landing gear is used, it

must be possible, during the take-off ground run

on concrete, to maintain any attitude up to thrust

line level, at 75% of VSR1 .

CS 25.233 Directional stability and

control

(a) There may be no uncontrollable groundlooping tendency in 90º cross winds, up to a wind

velocity of 37 km/h (20 kt) or 0·2 VSR0 , whichever is

greater, except that the wind velocity need not

exceed 46 km/h (25 kt) at any speed at which the

aeroplane may be expected to be operated on the

ground. This may be shown while establishing the

90º cross component of wind velocity required by CS

25.237.

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(b) Aeroplanes must be satisfactorily

controllable, without exceptional piloting skill or

alertness, in power-off landings at normal landing

speed, without using brakes or engine power to

maintain a straight path. This may be shown during

power-off landings made in conjunction with other

tests.

(c) The aeroplane must have adequate

directional control during taxying. This may be

shown during taxying prior to take-offs made in

conjunction with other tests.

CS 25.235 Taxying condition

The shock absorbing mechanism may not damage the

structure of the aeroplane when the aeroplane is

taxied on the roughest ground that may reasonably be

expected in normal operation.

CS 25.237 Wind velocities

(a) The following applies:

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(1) A 90º cross component of wind

velocity, demonstrated to be safe for take-off and

landing, must be established for dry runways and

must be at least 37 km/h (20 kt) or 0·2 VSR0,

whichever is greater, except that it need not

exceed 46 km/h (25 kt).

(2) The crosswind component for takeoff

established without ice accretions is valid in icing

conditions.

(3) The landing crosswind component

must be established for:

(i) Non-icing conditions, and

(ii) Icing conditions with the

landing ice accretion defined in appendix C.

[Amdt. No.:25/3]

MISCELLANEOUS FLIGHT REQUIREMENTS

CS 25.251 Vibration and buffeting

(a) The aeroplane must be demonstrated in

flight to be free from any vibration and buffeting that

would prevent continued safe flight in any likely

operating condition.

(b) Each part of the aeroplane must be

demonstrated in flight to be free from excessive

vibration under any appropriate speed and power

conditions up to VDF/MDF. The maximum speeds

shown must be used in establishing the operating

limitations of the aeroplane in accordance with CS

25.1505.

(c) Except as provided in sub-paragraph (d) of

this paragraph, there may be no buffeting condition,

in normal flight, including configuration changes

during cruise, severe enough to interfere with the

control of the aeroplane, to cause excessive fatigue to

the crew, or to cause structural damage. Stall warning

buffeting within these limits is allowable.

(d) There may be no perceptible buffeting

condition in the cruise configuration in straight flight

at any speed up to VMO/MMO, except that the stall

warning buffeting is allowable.

(e) For an aeroplane with MD greater than 0·6

or with a maximum operating altitude greater than

7620 m (25,000 ft), the positive manoeuvring load

factors at which the onset of perceptible buffeting

occurs must be determined with the aeroplane in the

cruise configuration for the ranges of airspeed or

Mach number, weight, and altitude for which the

aeroplane is to be certificated. The envelopes of load

factor, speed, altitude, and weight must provide a

sufficient range of speeds and load factors for normal

operations. Probable inadvertent excursions beyond

the boundaries of the buffet onset envelopes may not

result in unsafe conditions. (See AMC 25.251(e).)

[Amdt. No.:25/1]

CS 25.253 High-speed characteristics

(a) Speed increase and recovery

characteristics. The following speed increase and

recovery characteristics must be met:

(1) Operating conditions and characteristics likely to cause inadvertent speed increases

(including upsets in pitch and roll) must be

simulated with the aeroplane trimmed at any

likely cruise speed up to VMO/MMO. These

conditions and characteristics include gust upsets,

inadvertent control movements, low stick force

gradient in relation to control friction, passenger

movement, levelling off from climb, and descent

from Mach to air speed limit altitudes.

(2) Allowing for pilot reaction time after

effective inherent or artificial speed warning

occurs, it must be shown that the aeroplane can be

recovered to a normal attitude and its speed

reduced to VMO/MMO, without –

(i) Exceptional piloting strength or

skill;

(ii) Exceeding VD/MD, VDF/MDF, or

the structural limitations; and

(iii) Buffeting that would impair the

pilot’s ability to read the instruments or

control the aeroplane for recovery.

(3) With the aeroplane trimmed at any

speed up to VMO/MMO, there must be no reversal

of the response to control input about any axis at

any speed up to VDF/MDF. Any tendency to pitch,

roll, or yaw must be mild and readily controllable,

using normal piloting techniques. When the

aeroplane is trimmed at VMO/MMO, the slope of the

elevator control force versus speed curve need not

be stable at speeds greater than VFC/MFC, but there

must be a push force at all speeds up to VDF/MDF

and there must be no sudden or excessive

reduction of elevator control force as VDF/MDF is

reached.

(4) Adequate roll capability to assure a

prompt recovery from a lateral upset condition

must be available at any speed up to VDF/MDF.

(See AMC 25.253(a)(4).)

(5) Extension of speedbrakes. With the

aeroplane trimmed at VMO/MMO, extension of the

speedbrakes over the available range of

movements of the pilots control, at all speeds

above VMO/MMO, but not so high that VDF/MDF

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would be exceeded during the manoeuvre, must

not result in:

(i) An excessive positive load

factor when the pilot does not take action to

counteract the effects of extension;

(ii) Buffeting that would impair the

pilot’s ability to read the instruments or

control the aeroplane for recovery; or

(iii) A nose-down pitching moment,

unless it is small. (See AMC 25.253(a)(5).)

(6) Reserved

(b) Maximum speed for stability characteristics,

VFC/MFC. VFC/MFC is the maximum speed at which

the requirements of CS 25.143(g), 25.147(e),

25.175(b)(1), 25.177(a) through (c ), and 25.181

must be met with wing-flaps and landing gear

retracted. Except as noted in CS 25.253(c), VFC/MFC

may not be less than a speed midway between

VMO/MMO and VDF/MDF, except that, for altitudes

where Mach Number is the limiting factor, MFC need

not exceed the Mach Number at which effective

speed warning occurs.

(c) Maximum speed for stability characteristics

in icing conditions. The maximum speed for stability

characteristics with the ice accretions defined in

Appendix C, at which the requirements of CS

25.143(g), 25.147(e), 25.175(b)(1), 25.177(a)

through (c) and 25.181 must be met, is the lower of:

(1) 556 km/h (300 knots) CAS,

(2) VFC, or

(3) A speed at which it is demonstrated

that the airframe will be free of ice accretion due

to the effects of increased dynamic pressure.

[Amdt. No.:25/3]

CS 25.255 Out-of-trim characteristics

(See AMC 25.255)

(a) From an initial condition with the aeroplane

trimmed at cruise speeds up to VMO/MMO, the

aeroplane must have satisfactory manoeuvring

stability and controllability with the degree of out-oftrim in both the aeroplane nose-up and nose-down

directions, which results from the greater of –

(1) A three-second movement of the

longitudinal trim system at its normal rate for the

particular flight condition with no aerodynamic

load (or an equivalent degree of trim for

aeroplanes that do not have a power-operated trim

system), except as limited by stops in the trim

system, including those required by CS25.655 (b)

for adjustable stabilisers; or

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(2) The maximum mistrim that can be

sustained by the autopilot while maintaining level

flight in the high speed cruising condition.

(b) In the out-of-trim condition specified in

sub-paragraph (a) of this paragraph, when the normal

acceleration is varied from + 1 g to the positive and

negative values specified in sub-paragraph (c) of this

paragraph –

(1) The stick force vs. g curve must have

a positive slope at any speed up to and including

VFC/MFC; and

(2) At speeds between VFC/MFC and

VDF/MDF, the direction of the primary longitudinal

control force may not reverse.

(c) Except as provided in sub-paragraphs (d)

and (e) of this paragraph compliance with the

provisions of sub-paragraph (a) of this paragraph

must be demonstrated in flight over the acceleration

range –

(1) –1g to 2·5 g; or

(2) 0 g to 2·0 g, and extrapolating by an

acceptable method to – 1 g and 2·5 g.

(d) If the procedure set forth in sub-paragraph

(c)(2) of this paragraph is used to demonstrate

compliance and marginal conditions exist during

flight test with regard to reversal of primary

longitudinal control force, flight tests must be

accomplished from the normal acceleration at which

a marginal condition is found to exist to the

applicable limit specified in sub-paragraph (c)(1) of

this paragraph.

(e) During flight tests required by subparagraph (a) of this paragraph the limit manoeuvring

load factors prescribed in CS25.333 (b) and 25.337,

and the manoeuvring load factors associated with

probable inadvertent excursions beyond the

boundaries of the buffet onset envelopes determined

under CS 25.251 (e), need not be exceeded. In

addition, the entry speeds for flight test

demonstrations at normal acceleration values less

than 1 g must be limited to the extent necessary to

accomplish a recovery without exceeding VDF/MDF.

(f) In the out-of-trim condition specified in

sub-paragraph (a) of this paragraph, it must be

possible from an overspeed condition at VDF/MDF, to

produce at least 1·5 g for recovery by applying not

more than 556 N (125 lbf) of longitudinal control

force using either the primary longitudinal control

alone or the primary longitudinal control and the

longitudinal trim system. If the longitudinal trim is

used to assist in producing the required load factor, it

must be shown at VDF/MDF that the longitudinal trim

can be actuated in the aeroplane nose-up direction

with the primary surface loaded to correspond to the

1-B-24

Annex to ED Decision 2008/006/R

Amendment 5

CS-25 BOOK 1

least of the following aeroplane nose-up control

forces:

(1) The maximum control forces expected

in service as specified in CS 25.301 and 25.397.

(2) The control force required to produce

1·5 g.

(3) The control force corresponding to

buffeting or other phenomena of such intensity

that it is a strong deterrent to further application

of primary longitudinal control force.

1-B-25

Annex to ED Decision 2008/006/R

Amendment 5

CS–25 BOOK 1

GENERAL

CS 25.301 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. (See AMC No. 1 to CS

25.301(b).) 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. (See AMC No. 2 to CS 25.301(b).)

(c) If deflections under load would significantly

change the distribution of external or internal loads,

this redistribution must be taken into account.

[Amdt. No.:25/1]

CS 25.302 Interaction of systems and

structures

For aeroplanes equipped with systems that affect

structural performance, either directly or as a result

of a failure or malfunction, the influence of these

systems and their failure conditions must be taken

into account when showing compliance with the

requirements of Subparts C and D. Appendix K of

CS-25 must be used to evaluate the structural

performance of aeroplanes equipped with these

systems.

[Amdt. No.:25/1]

CS 25.303 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

loading condition is prescribed in terms of ultimate

loads, a factor of safety need not be applied unless

otherwise specified.

帅哥

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

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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.

帅哥

(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.

帅哥

(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;

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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

帅哥

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:

⎥