currently with the airplane load factor of 1.0 (points A1 to D1, § 25.333(b)). This positive acceleration must be equal 39 to at least n (n-1.5) (radians/sec.2), V where

(a) n is the positive load factor at the speed under consideration; and

(b) V is the airplane equivalent speed in knots.

(ii) Unless lesser values could not be exceeded, a negative pitching acceleration (nose down) is assumed to be reached concurrently with the positive maneuvering load factor (points A, to D2, § 25.333 (b)). This negative pitching acceleration must be equal to at least

sult in pitching accelerations not less than those specified in subparagraph (2). (d) Gust conditions. The gust conditions B' through J', § 25.333 (c), must be investigated. The following provisions apply:

(1) The air load increment due to a specified gust must be added to the initial balancing tail load corresponding to steady level flight.

(2) The alleviating effect of wing down-wash and of the airplane's motion in response to the gust may be included in computing the tail gust load increment.

(3) Instead of a rational investigation of the airplane response, the gust alleviation factor K, may be applied to the specified gust intensity for the horizontal tail.

§ 25.333 Flight 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 maneuvering and gust envelopes (V-n diagrams) of paragraphs (b) and (c) of this section. These envelopes must also be used in determining the airplane structural operating limitations as specified in § 25.1501.

(b) Maneuvering envelope.

§ 25.335 Design airspeeds.

The selected design airspeeds are equivalent airspeeds (EAS). Estimated values of V and V must be conserva$1 tive.

(a) Design cruising speed, Vc For Vo, the following apply:

C

(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) In the absence of a rational investigation substantiating the use of other values, Vc may not be less than VB+43 knots. However, it 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, Ve may be limited to a selected Mach number.

(b) Design dive speed, VD. The selected design dive speed must be used in determining the maximum operating limit speed for the airplane in accordance with § 25.1505.

(c) Design maneuvering speed VA· For VA, the following apply:

(ii) V is the stalling speed with flaps 81 retracted.

(2) VA and Vs must be evaluated at the design weight and altitude under consideration.

(3) VA need not be more than Vo or the speed at which the positive CN mas curve intersects the positive maneuver load factor line, whichever is less.

(d) Design speed for maximum gust intensity, VB For VB, the following apply:

B

B.

(1) V, may not be less than the speed determined by the intersection of the line representing the maximum positive lift CN max and the line representing the rough air gust velocity on the gust V-n diagram, or (Vn) Vs1, whichever is less, where

(i) ng is the positive airplane gust load factor due to gust, at speed Vo (in accordance with § 25.341), and at the particular weight under consideration; and

(ii) V is the stalling speed with the $1 flaps retracted at the particular weight under consideration.

B

(2) VR need not be greater than Vc. (e) Design flap speeds, VF. For VF

the following apply:

(1) The design flap speed for each flap position (established in accordance with § 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 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 may not be less than—

(i) 1.6 Vs, with the flaps in takeoff position at maximum takeoff weight; (ii) 1.8 Vs, with the flaps in approach position at maximum landing weight; and

(iii) 1.8 Vg with the flaps in landing So position at maximum landing weight. § 25.337

tors.

Limit maneuvering load fac

(a) Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the limit maneuvering load factors prescribed in this section. Pitching velocities appropriate to the corresponding pull-up and steady turn maneuvers must be taken into account.

(b) The positive limit maneuvering load factor n for any speed up to V, may not be less than 2.5.

(c) The negative limit maneuvering 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 Vo to zero at VD.

(d) Maneuvering load factors lower than those specified in this section may be used if the airplane has design features that make it impossible to exceed these values in flight.

Cag

airplane mass ratio;

Ude derived gust velocities referred to in paragraph (a) (fps);

p density of air (slugs/cu. ft.); W/Swing loading (psf);

C= mean geometric chord (ft.);
g=acceleration due to gravity (ft/
sec.2);

V airplane equivalent speed (knots);
and
a=slope of the airplane normal force
coefficient curve CNA per radian
if the gust loads are applied to the
wings and horizontal tail surfaces
simultaneously
by a rational
method. The wing lift curve slope
CL per radian may be used when
the gust load is applied to the
wings only and the horizontal
tail gust loads are treated as a
separate condition.

§ 25.343

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 § 25.1001 (c), 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 subpart.

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 maneuvering load factor of +2.25; and

(ii) Gust intensities equal to 85 percent of the values prescribed in § 25.341; and

(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.

§ 25.345 High lift devices.

(a) If flaps are to be used during takeoff, approach, or landing, at the design flap speeds established for these stages of flight under § 25.335(e) and with the flaps in the corresponding positions, the airplane is assumed to be subjected to symmetrical maneuvers and gusts within the range determined by

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

(2) Positive and negative 25 fps derived gusts acting normal to the flight path in level flight.

(b) The airplane must be designed for the conditions prescribed in paragraph (a) of this section, except that the airplane 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, 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 head-on gust of 25 feet per second velocity (EAS).

(c) If flaps or similar high lift devices are to be used as speed brakes in en route conditions, and with flaps in the appropriate position at speeds up to the flap design speed chosen for these conditions,

The airplane must be designed for rolling loads resulting from the conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces.

(a) Maneuvering. The following conditions, speeds, and aileron deflections (except as the deflections may be limited by pilot effort) must be considered in combination with an airplane load factor of zero and of two-thirds of the positive maneuvering factor used in design. In determining the required aileron deflections, the torsional flexibility of the wing must be considered in accordance with § 25.301(b):

(1) Conditions corresponding to steady rolling velocities must be investigated. In addition, conditions corresponding to maximum angular acceleration must be investigated for airplanes 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 ma

D'

(4) At V the aileron deflection must be that required to produce a rate of roll not less than one-third of that in subparagraph (2) of this paragraph.

(b) Unsymmetrical gusts. The condition of unsymmetrical gusts must be considered by modifying the symmetrical flight conditions B' or C' (in § 25.333 (c)) whichever produces the greater load factor. It is assumed that 100 percent of the wing air load acts on one side of the airplane and 80 percent acts on the other side.

§ 25.351 Yawing conditions.

The airplane must be designed for loads resulting from the conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the principal masses furnishing the reacting inertia forces: (a) Maneuvering. At speeds from V MC to VA, the following maneuvers must be considered. In computing the tail loads, the yawing velocity may be assumed to be zero:

(1) With the airplane in unaccelerated flight at zero yaw, it is assumed that the rudder control is suddenly displaced to the maximum deflection, as limited by the control stops or by a 300 lb. rudder pedal force, whichever is critical.

(2) With the rudder deflected as specified in subparagraph (1) of this paragraph, it is assumed that the airplane yaws to the resulting sideslip angle.

(3) With the airplane yawed to the static sideslip angle corresponding to the rudder deflection specified in subparagraph (1) of this paragraph, it is assumed that the rudder is returned to neutral.

(b) Lateral gusts. The airplane is assumed to encounter derived gusts normal to the plane of symmetry while in unaccelerated flight. The derived gusts and airplane speeds corresponding to conditions B' through J' (in § 25.333 (c)) (as determined by §§ 25.341 and 25.345 (a) (2) or 25.345 (c) (2)) must be investigated. The shape of the gust must be as specified in § 25.341. In the absence of a rational investigation of the airplane's response to a gust, the gust loading on the vertical tail surfaces must be computed as follows:

V airplane equivalent speed (knots).
SUPPLEMENTARY CONDITIONS

§ 25.361 Engine torque.

(a) Each engine mount and its supporting structures must be designed for engine torque effects combined with—

(1) The limit engine torque corresponding to takeoff power and propeller speed acting simultaneously with 75 percent of the limit loads from flight condition A of § 25.333 (b);

(2) The limit engine torque corresponding to maximum continuous power and propeller speed acting simultaneously with the limit loads from flight condition A of § 25.333 (b); and

(3) For turbopropeller installations, in addition to the conditions specified in subparagraphs (1) and (2) of this paragraph, the limit engine torque corresponding to takeoff power and propeller speed multiplied by a factor of 1.6 acting simultaneously with 1g level flight loads.

(b) For turbine engine installations, the limit engine torque load imposed by sudden engine stoppage due to malfunction or structural failure (such as compressor jamming) must be considered in the design of the engine mounts and supporting structure.

(c) The limit engine torque is obtained by multiplying the mean torque for maximum continuous power by a factor of

(1) 1.25 for turbopropeller installations;

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

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

§ 25.363 Side load on engine mount.

(a) Each engine mount and its supporting structure must be designed for a limit load factor in a lateral direction, for the side load on the engine mount, at least equal to the maximum load factor obtained in the yawing conditions but not less than

(1) 1.33; or