nertia loads corresponding to any of the symmetrical flight conditions specified n§§ 23.331 through 23.341.

(b) The incremental horizontal tail loads due to maneuvering and gusts must be reacted by the angular inertia of the airplane in a rational or conservative

manner.

$23.333 Flight envelope.

(a) General. Compliance with the strength requirements of this subpart must be shown at any combination of airspeed and load factor on and within the boundaries of a flight envelope (similar to the one in paragraph (d) of this section) that represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria of paragraphs (b) and (c) of this section respectively.

(b) Maneuvering envelope. Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the following limit load factors:

(1) The positive maneuvering load factor specified in § 23.337 at speeds up to Vpi

(2) The negative maneuvering load factor specified in § 23.337 at Vo; and (3) Factors varying linearly with speed from the specified value at Vo to

0.0 at VD for the normal category, and -1.0 at VD for the acrobatic and utility categories.

(c) Gust envelope. (1) The airplane is assumed to be subjected to symmetrical vertical gusts in level flight. The resulting limit load factors must correspond to the conditions determined as follows:

(i) Positive (up) and negative (down) gusts of 50 f.p.s. at Vc must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 50 f.p.s. at 20,000 feet to 25 f.p.s. at 50,000 feet.

(ii) Positive and negative gusts of 25 f.p.s. at VD must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 25 f.p.s. at 20,000 feet to 12.5 f.p.s. at 50,000 feet.

(2) The following assumptions must be made:

(i) The shape of the gust is

Except as provided in paragraph (a) (4) of this section, the selected design airspeeds are equivalent airspeeds (EAS).

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

(1) Vc (in knots) may not be less than

(i) 33\/W/S (for normal and utility category airplanes); and

(ii) 36W/S (for acrobatic category airplanes).

(2) For values of W/S more than 20, the multiplying factors may be decreased linearly with W/S to a value of 28.6 where W/S=100.

(3) V. need not be more than 0.9 VH at sea level.

(4) At altitudes where an MD is established, a cruising speed Mc limited by compressibility may be selected.

(b) Design dive speed VD. For VD, the following apply:

(1) VD/MD may not be less than 1.25 Vc/Mc; and

(2) With Vo min, the required minimum design cruising speed, VD (in knots) may not be less than

(i) 1.40 Vc min (for normal category airplanes);

(ii) 1.50 Vo min (for utility category airplanes); and

(iii) 1.55 Vc min (for acrobatic category airplanes).

(3) For values of W/S more than 20, the multiplying factors in subparagraph (2) of this paragraph may be decreased linearly with W/S to a value of 1.35 where W/S=100.

(4) Compliance with subparagraphs (1) and (2) of this paragraph need not be shown if VD/MD is selected so that the minimum speed margin between Vo/Mc and VD/MD is the greater of the following:

(i) The speed increase resulting when, from the initial condition of stabilized flight at Ve/Mc, the airplane is assumed to be upset, flown for 20 seconds along a flight path 7.5° below the initial path, and then pulled up with a load factor of 1.5 (0.5 g. acceleration increment). At least 75 percent maximum continuous power for reciprocating engines, and maximum cruising power for turbines, or, if less, the power required for Vo/Mc for both kinds of engines, must be assumed until the pullup is initiated, at which point power reduction and pilotcontrolled drag devices may be used.

(ii) Mach 0.05 (at altitudes where a MD is established).

(c) Design maneuvering speed V For V, the following applies:

A

(1) V (in miles per hour) may not b less than V√n where—

(i) Vg is a computed stalling spee with flaps retracted at the design weight normally based on the maximum air plane normal force coefficients, CN; an

(ii) n is the limit maneuvering loa factor used in design.

(2) The value of VA need not excee the value of Vo used in design.

[Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964 as amended by Amdt. No. 23-7, 34 F.R. 13088 Aug. 13, 1969]

(3) 6.0 for acrobatic category airplanes.

(b) The negative limit maneuvering load factor may not be less than

(1) 0.4 times the positive load factor for the normal and utility categories; or (2) 0.5 times the positive load factor for the acrobatic category.

(c) Maneuvering load factors lowe than those specified in this section may be used if the airplane has design fea tures that make it impossible to excee these values in flight.

[Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964 as amended by Amdt. No. 23-7, 34 F.R. 13088 Aug. 13, 1969]

§ 23.341 Gust loads factors.

In the absence of a more rationa analysis, the gust load factors must be computed as follows:

W/S Wing loading (p.s.f.); 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 С 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.

Amdt. No. 23-7, 34 F.R. 13088, Aug. 13, 1969] 23.345 High lift devices.

(a) If flaps or similar high lift devices be used for takeoff, approach, or landg are installed, the airplane, with the aps fully deflected at V, is assumed to e subjected to symmetrical maneuvers nd gusts resulting in limit load factors within the range determined by

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

(2) Positive and negative gust of 25 eet per second acting normal to the Hight path in level flight.

(b) V, must be assumed to be not less han 1.4 Vg or 1.8 Vgr. whichever is greater, where

Vs is the computed stalling speed with flaps retracted at the design weight; and VSF is the computed stalling speed with flaps fully extended at the design weight. However, if an automatic flap load limting device is used, the airplane may be lesigned for the critical combinations of irspeed and flap position allowed by hat device.

(c) In designing the flaps and suporting structures, slipstream effects nust be accounted for, as specified in aragraph (b) of § 23.457.

(d) In determining external loads on he airplane as a whole, thrust, slipstream, and pitching acceleration may be assumed to be zero.

(e) The requirements of § 23.457, and this section may be complied with separately or in combination.

Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964, as amended by Amdt. No. 23-7, 34 F.R. 13088, Aug. 13, 1969]

$23.347 Unsymmetrical flight condi

tions.

The airplane is assumed to be subJected to the unsymmetrical flight con

The wing and wing bracing must be designed for the following loading conditions:

(a) Unsymmetrical wing loads appropriate to the category. Unless the following values result in unrealistic loads, the rolling accelerations may be obtained by modifying the symmetrical flight conditions in § 23.333 (d) as follows:

(1) For the acrobatic category, in conditions A and F, assume that 100 percent of the semispan wing airload acts on one side of the plane of symmetry and 60 percent of this load acts on the other side.

(2) For the normal and utility categories, in condition A, assume that 100 percent of the semispan wing airload acts on one side of the airplane and 70 percent of this load acts on the other side. For airplanes of more than 1,000 pounds design weight, the latter percentage may be increased linearly with weight up to 75 percent at 12,500 pounds.

(b) The loads resulting from the alleron deflections and speeds specified in § 23.455, in combination with an airplane load factor of at least two thirds of the positive maneuvering load factor used for design. Unless the following values result in unrealistic loads, the effect of aileron displacement on wing torsion may be accounted for by adding the following increment to the basic airfoil moment coefficient over the aileron portion of the span in the critical condition determined in § 23.333 (d):

§ 23.361 Engine torque.

(a) Each engine mount and its supporting structure must be designed for the effects of

(1) The limit torque corresponding to takeoff power and propeller speed acting simultaneously with 75 percent of the limit loads from flight condition A of § 23.333 (d);

(2) The limit torque corresponding to the maximum continuous power and propeller speed, acting simultaneously with the limit loads from flight condition A of § 23.333 (d); and

(3) For turbopropeller installations, in addition to the condition specified in subparagraphs (1) and (2) of this paragraph, the limit engine torque corresponding to takeoff power and propeller speed, multiplied by a factor accounting for propeller control system malfunction, including quick feathering, acting simultaneously with 1g. level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used.

(b) The limit torque is obtained by multiplying the mean torque by a factor of

(1) 1.25 for turbopropeller installations;

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

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

(c) Engine torque effects need not be investigated for any other conditions. [Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964, as amended by Amdt. No. 23-7, 34 F.R. 13088, Aug. 13, 1969]

§ 23.363 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, of not less than

loads combined with pressure differentia loads from zero up to the maximu relief valve setting.

(b) The external pressure distributio in flight, and any stress concentration must be accounted for.

(c) If landings may be made, with th cabin pressurized, landing loads must b combined with pressure differenti loads from zero up to the maximum a lowed during landing.

(d) The airplane structure must b strong enough to withstand the pressu differential loads corresponding to th maximum relief valve setting multiplie by a factor of 1.33, omitting other load

(e) If a pressurized cabin has two more compartments separated by bull heads or a floor, the primary structu must be designed for the effects of su den release of pressure in any compar ment with external doors or window This condition must be investigated f the effects of failure of the largest oper ing in the compartment. The effects intercompartmental venting may be con

sidered.

§ 23.367

Unsymmetrical loads due engine failure.

(a) Tubopropeller airplanes must b designed for the unsymmetrical load resulting from the failure of the critica engine including the following condition in combination with a single malfunction of the propeller drag limiting sys tem, considering the probable pilot cor rective action on the flight controls:

(1) At speeds between Vс and V the loads resulting from power failur because of fuel flow interruption are con sidered to be limit loads.

(2) At speeds between VMC and Vc, th loads resulting from the disconnectio of the engine compressor from the tu bine or from loss of the turbine blad are considered to be ultimate loads.

(3) The time history of the thru decay and drag buildup occurring as result of the prescribed engine failure must be substantiated by test or othe data applicable to the particular engine propeller combination.

(4) The timing and magnitude of th probable pilot corrective action must b conservatively estimated, considering the characteristics of the particular engine propeller-airplane combination.

(b) Pilot corrective action may be as sumed to be initiated at the time maxi mum yawing velocity is reached, but no

ier than 2 seconds after the engine ure. The magnitude of the corrective on may be based on the limit pilot ces specified in § 23.397 except that er forces may be assumed where it is wn by analysis or test that these ces can control the yaw and roll reting from the prescribed engine failconditions.

ndt. No. 23-7, 34 F.R. 13089, Aug. 13, 1969] 3.369 Special conditions for rear lift

truss.

a) If a rear lift truss is used, it must designed for conditions of reversed flow at a design speed of―

1-8.7/W/S+8.7(knots)

(b) Either aerodynamic data for the rticular wing section used, or a value C equalling -0.8 with a chordwise tribution that is triangular between a ak at the trailing edge and zero at e leading edge, must be used.

ocket No. 4080, 29 F.R. 17955, Dec. 18, 1964, amended by Amdt. No. 23-7, 34 F.R. 13089, g. 13, 1969; 34 F.R. 17509, Oct. 30, 1969]

23.371 Gyroscopic loads.

For turbopropeller powered airplanes, ch engine mount and its supporting ructure must be designed for the gyroopic loads that result, with the engines t maximum continuous r.p.m., under ither of the following conditions: (a) The

conditions prescribed in

§ 23.351 and 23.423.

(b) All possible combinations of the ›llowing:

(1) A yaw velocity of 2.5 radians per cond.

(2) A pitch velocity of 1 radian per Cond.

(3) A normal load factor of 2.5. (4) Maximum continuous thrust. ndt. No. 23-7, 34 F.R. 13089, Aug. 13, 1969]

3.373 Speed control devices.

If speed control devices (such as spoiland drag flaps) are incorporated for ein enroute conditions(a) The airplane must be designed r the symmetrical maneuvers and gusts rescribed in §§ 23.333, 23.337, and 3.341, and the yawing maneuvers and ateral gusts in §§ 23.441 and 23.443, with the device extended at speeds up to the placard device extended speed; and (b) If the device has automatic operting or load limiting features, the airlane must be designed for the maneuver

and gust conditions prescribed in paragraph (a) of this section at the speeds and corresponding device positions that the mechanism allows.

[Amdt. No. 23-7, 34 F.R. 13089, Aug. 13, 1969] CONTROL SURFACE AND SYSTEM LOADS

§ 23.391 Control surface loads.

(a) The control surface loads specified in §§ 23.397 through 23.459 are assumed to occur in the conditions described in §§ 23.331 through 23.351.

(b) If allowed by the following sections, the values of control surface loading in Appendix B of this part may be used, instead of particular control surface data, to determine the detailed rational requirements of §§ 23.397 through 23.459, unless these values result in unrealistic loads.

§ 23.395 Control system.

(a) Each flight control system and its supporting structure must be designed for loads corresponding to at least 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in §§ 23.391 through 23.459. In addition, the following apply:

(1) The system limit loads need not exceed the higher of the loads that can be produced by the pilot and automatic devices operating the controls. However, autopilot forces need not be added to pilot forces. The system must be designed for the maximum effort of the pilot or autopilot, whichever is higher. In addition, if the pilot and the autopilot act in opposition, the part of the system between them may be designed for the maximum effort of the one that imposes the lesser load. Pilot forces used for design need not exceed the maximum forces prescribed in § 23.397(b).

(2) The design must, in any case, provide a rugged system for service use, considering jamming, ground gusts, taxiing downwind, control inertia, and friction. Compliance with this subparagraph may be shown by designing for loads resulting from application of the minimum forces prescribed in § 23.397 (b).

(b) A 125 percent factor on computed hinge moments must be used to design elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be used if hinge moments are based on ac