Section 1 – requirements

The rotorcraft must be designed for –

(a) A limit manoeuvring load factor ranging

from a positive limit of 3.5 to a negative limit of

– 1.0; or

(b) Any positive limit manoeuvring load factor

not less than 2.0 and any negative limit manoeuvring

load factor of not less than – 0.5 for which –

(1) The probability of being exceeded

is shown by analysis and flight tests to be

extremely remote; and

(2) The selected values are appropriate

to each weight condition between the design

maximum and design minimum weights.

The loads resulting from the application of

limit manoeuvring load factors are assumed to act

at the centre of each rotor hub and at each

auxiliary lifting surface, and to act in directions,

and with distributions of load among the rotors

and auxiliary lifting surfaces, so as to represent

each critical manoeuvring condition, including

power-on and power-off flight with the maximum

design rotor tip speed ratio. The rotor tip speed

ratio is the ratio of the rotorcraft flight velocity

component in the plane of the rotor disc to the

rotational tip speed of the rotor blades, and is

expressed as follows:

µ = ------------

ΩR

where –

a = The angle between the projection, in the

plane of symmetry, of the axis of no

feathering and a line perpendicular to the

flight path (radians, positive when the axis

is pointing aft);

_ = The angular velocity of rotor (radians per

second); and

R = The rotor radius (ft).

µ = ------------

ΩR

където:

а = ъгълът между проекцията, в равнината на симетрия, на оста без флюгиране и линията, перпендикулярна на траекторията на полета (радиани, положителни, когато оста сочи към опашката на ророркрафта);

Ω = ъгловата скорост на носещия винт (радиани в секунда); и

R = радиусът на носещия винт (фута).

The rotorcraft must be designed to withstand,

at each critical airspeed including hovering, the

loads resulting from a vertical gust of 9.1 metres

per second (30 ft/s).

(a) Each rotorcaft must be designed for the

loads resulting from the manoeuvres specified in

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

with –

(1) Unbalanced aerodynamic moments

reacts to in a rational or conservative manner

considering the principal masses furnishing the

reacting inertia forces; and

(2) Maximum main rotor speed.

(b) To produce the load required in subparagraph

(a) of this paragraph, in unaccelerated

flight with zero yaw, at forward speeds from zero

up to 0.6 VNE –

(1) Displace the cockpit directional

control suddenly to the maximum deflection

limited by the control stops or by the

(a);

90°, whichever is less; and

(3) Return the directional control

suddenly to neutral.

(c) To produce the load required in subparagraph

(a) of this paragraph, in unaccelerated

flight with zero yaw, at forward speeds from

0.6 VNE up to VNE or VH, whichever is less –

(1) Displace the cockpit directional

control suddenly to the maximum deflection

limited by the control stops or by the

[maximum pilot force specified in JAR 27.397

(a); ]

(2) Attain a resulting sideslip angle or

15°, whichever is less, at the lesser speed of

VNE or VH;

(3) Vary the sideslip angles of subparagraphs

(b) (2) and (c) (2) of this paragraph

directly with speed; and

(4) Return the directional control

suddenly to neutral.

(2) Максималната скорост на главния носещ винт.

(b) Да създаде натоварването, което се изисква в подпараграф (а) на настоящия параграф, при установен полет със нулево отклонение от курса при постъпателни скорости от нула до 0.6 V_NE.

(1) Да отклони енергично попътното управление от пилотската кабина до максималното отклонение , ограничавано от ограничителите за управление или от максималното усилие на пилота, посочено конкретно в JAR 27.397 (а);

(2) Да постигне резултатен ъгъл на странично плъзгане или 90 градуса, което е по-малко; и

(3) Да върне попътното управление енергично на нула/в неутрално положение.

(с) Да създаде натоварването, което се изисква в подпараграф (а) на настоящия параграф, при установен полет със нулево отклонение от курса при постъпателни скорости от 0.6 V_NE до V_NE или Vн, която е по-малка-

(1) Да отклони енергично попътното управление от пилотската кабина до максималното отклонение , ограничавано от ограничителите за управление или от максималното усилие на пилота, посочено конкретно в JAR 27.397 (а);]

(2) Да постигне резултатен ъгъл на странично плъзгане или 15 градуса, което е по-малко, при по-малката скорост V_NE или Vн ;

3. 
   Да променя ъглите на странично плъзгане, посочени в подпараграфи (b) 2 е (с) (2) на настоящия параграф директно със скоростта; и
   
4. 
   Да върне попътното управление енергично на нула/в неутрално положение.
   

(a) For turbine engines, the limit torque may

not be less than the highest of –

(1) The mean torque for maximum

continuous power multiplied by 1.25;

(2) The torque required by JAR 27.923;

(3) The torque required by JAR 27.927;

or

(4) The torque imposed by sudden

structural failure (such as compressor

jamming).

(b) For reciprocating engines, the limit

torque may not be less than the mean torque for

maximum continuous power multiplied by –

(1) 1.33, for engines with five or more

cylinders; and

(2) Two, three, and four, for engines

with four, three, and two cylinders,

respectively.

]

(а) За турбинни двигатели, -

граничния въртящ момент на вала на двигателя не трябва да е по-малък от :

1. 
   Средния въртящ момент на вала за максимална продължителна мощност, умножено с 1.25;
   
2. 
   Въртящият момент на вала на двигателя, коийто се изисква от JAR 27.923;
   
3. 
   Въртящият момент на вала на двигателя, коийто се изисква от JAR 27.927; и
   
4. 
   Въртящият момент на вала на двигателя, който принудително се създава при внезапното спиране на двигателя поради авария или разрушаване на конструкцията (като например заклинване на компресора).
   

1. 
   1.33 за двигатели с пет или повече цилиндри; и
   
2. 
   две, три и четири за двигатели , съответно с четири, три и два цилиндъра.
   

Each auxiliary rotor, each fixed or movable

stabilising or control surface, and each system

operating any flight control must meet the

requirements of JAR 27.395, 27.397, 27.399,

27.411 and 27.427.

(a) The part of each control system from the

pilot’s controls to the control stops must be designed

to withstand pilot forces of not less than –

(1) The forces specified in JAR 27.397;

or

(2) If the system prevents the pilot from

the maximum forces that the system allows the

pilot to apply, but not less than 0.60 times the

forces specified in JAR 27.397.

(b) Each primary control system including its

supporting structure, must be designed as follows:

(1) The system must withstand loads

resulting from the limit pilot forces prescribed

in JAR 27.397.

(2) Notwithstanding sub-paragraph (b)(3)

of this paragraph, when power-operated

actuator controls or power boost controls are

used, the system must also withstand the loads

resulting from the force output of each

normally energised power device, including

any single power boost or actuator system

failure.

(3) If the system design or the normal

operating loads are such that a part of the

system cannot react to the limit forces

prescribed in JAR 27.397, that part of the

system must be designed to withstand the

maximum loads that can be obtained in normal

operation. The minimum design loads must, in

any case, provide a rugged system for service

use, including consideration of fatigue,

jamming, ground gusts, control inertia and

friction loads. In the absence of rational

analysis, the design loads resulting from 0.60

of the specified limit pilot forces are

acceptable minimum design loads.

(4) If operational loads may be

exceeded through jamming, ground gusts,

control inertia, or friction, the system must

withstand the limit pilot forces specified in

JAR 27.397, without yielding.

(a) Except as provided in sub-paragraph (b)

of this paragraph the limit pilot forces are as

follows:

(130 pounds).

(100 pounds) fore and aft, and 298 Newtons]

(67 pounds) laterally.

(b) For flap, tab, stabiliser, rotor brake, and

landing gear operating controls, the following

[apply:

(1) Crank, wheel, and lever controls,

(25.4 + R) x 2.919 Newtons, where R = radius

in millimetres ( 







R 1 x 50 pounds, where R =

radius in inches), but not less than

222 Newtons (50 pounds) nor more than

445 Newtons (100 pounds) for hand-operated

controls or 578 Newtons (130 pounds) for footoperated

controls, applied at any angle within

20° of the plane of motion of the control.

(2) Twist controls, 356 x R Newtonmillimetres,

where R = radius in millimetres

(80 x R inch-pounds where R = radius in inches).

Освен както е предвидено в подпараграф (b) на настоящия параграф, максималните усилия на пилота са както следва:

[(1) За педалите за управление, 578 нютона] (130 фунта).

[(2) За лостовете за управление, 445 нютона (100 фунта) напред и назад, и 298 нютона] (67 фунта) напречно.

(b) За експлоатационните управляващи устройства за елерони, тримери, стабилизатора, спирачката на носещия вал и колесниците, [се прилага следното:

(1) Ръчка, кормило и лост (25.4 + R) х 2.919 нютона, където R е радиусът в милиметри ( 1 + R /3 х 50 фунта , където R е радиусът в инчове), но не по-малко от 222 нютона (50 фунта) нито повече от 445 нютона (100) фунта за ръчно задействани устройства за управление или 578 нютона (130 фунта) за крачно задайствани устройства за управление, прилагани при всякакъв ъгъл до20⁰ на плоскоста на движение на устройството за управление.

(2) Управляващи устройства със завъртане, 356 х R нютон –милиметри, където R = радиусът в милиметри (80 х R инч-фунта, където R = радиусът в инчове)].

Each dual primary flight control system must

be designed to withstand the loads that result

when pilot forces of 0.75 times those obtained

(a) In opposition; and

(b) In the same direction.

(a) It must be impossible for the tail rotor to

contact the landing surface during a normal

landing.

compliance with sub-paragraph (a) of this

paragraph –

(1) Suitable design loads must be

established for the guard; and

(2) The guard and its supporting

structure must be designed to withstand those

loads.

(b) Ако е необходим предпазител за опашния винт, за да се покаже съответствие с подпараграф (а) на този параграф:

1. 
   Трябва да се определят подходящи проектни натоварвания за предпазителя; и
   
2. 
   Предпазителят и неговата подпорна конструкция трябва да са проектирани да издържат тези натоварвания.
   

(a) Horizontal tail surfaces and their

supporting structure must be designed for

unsymmetrical loads arising from yawing and

rotor wake effects in combination with the

prescribed flight conditions.

(b) To meet the design criteria of subparagraph

(a) of this paragraph, in the absence of

more rational data, both of the following must be

met:

the symmetrical flight conditions acts on the

surface on one side of the plane of symmetry

and no loading acts on the other side.

(2) 50% of the maximum loading from

the symmetrical flight conditions acts on the

surface on each side of the plane of symmetry

but in opposite directions.

(c) For empennage arrangements where the

horizontal tail surfaces are supported by the

vertical tail surfaces, the vertical tail surfaces and

supporting structure must be designed for the

resulting from each prescribed flight condition,

considered separately. The flight conditions must

be selected so the maximum design loads are

obtained on each surface. In the absence of more

rational data, the unsymmetrical horizontal tail

surface loading distributions described in this

paragraph must be assumed

(a) Loads and equilibrium. For limit ground

loads –

(1) The limit ground loads obtained in

considered to be external loads that would

occur in the rotorcraft structure if it were

acting as a rigid body; and

(2) In each specified landing condition,

the external loads must be placed in

equilibrium with 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.