b. The loads reacted by airplane inertia.

§ 4b.257 Auxiliary float loads.

Auxiliary floats, their attachments, and supporting structure shall be designed for the following conditions. In the cases specified in paragraphs (a), (b), (c), and (d) of this section it shall be acceptable to distribute the prescribed water loads over the float bottom to avoid excessive local loads, using bottom pressures not less than those prescribed in paragraph (f) of this section.

(a) Step loading. The resultant water load shall be applied in the plane of symmetry of the float at a point three-fourths of the distance from the bow to the step and shall be perpendicular to the keel. The resultant limit load shall be computed as follows, except that the value of L need not exceed three times the weight of the displaced water when the float is completely submerged: C. V2 W2/3

where:

L=

tan2/38, (1+73)2/3

r, ratio of the lateral distance between the center of gravity and the plane of symmetry of the float to the radius of gyration in roll.

(b) Bow loading. The resultant limit load shall be applied in the plane of symmetry of the float at a point onefourth of the distance from the bow to the step and shall be perpendicular to the tangent to the keel line at that point. The magnitude of the resultant load shall be that specified in paragraph (a) of this section.

(c) Unsymmetrical step loading. The resultant water load shall consist of a component equal to 0.75 times the load specified in paragraph (a) of this section and a side component equal to 0.25 tan 8 times the load specified in paragraph (a) of this section. The side load shall be applied perpendicularly to the plane of symmetry of the float at a point midway between the keel and the chine.

(d) Unsymmetrical bow loading. The resultant water load shall consist of a component equal to 0.75 times the load specified in paragraph (b) of this section and a side component equal to 0.25 tan 8 times the load specified in paragraph (b) of this section. The side load shall be applied perpendicularly to the plane of symmetry at a point midway between the keel and the chine.

(e) Immersed float condition. The resultant load shall be applied at the centroid of the cross section of the float at a point one-third of the distance from

The following requirements deal with emergency conditions of landing on land or water in which the safety of the occupants shall be considered, although it is accepted that parts of the airplane may be damaged.

(a) The structure shall be designed to give every reasonable probability that all of the occupants, if they make proper use of the seats, belts, and other provisions made in the design (see § 4b.358), will escape serious injury in the event of a minor crash landing (with wheels up if the airplane is equipped with retractable landing gear) in which the occupants experience the following ultimate inertia forces relative to the surrounding structure:

agraph (a) of this section shall be acceptable if it is shown that the airplane structure can absorb the landing loads corresponding with the design landing weight and an ultimate descent velocity of 5 f. p. s. without exceeding the value chosen.

(c) The inertia forces specified in paragraph (a) of this section shall be applied to all items of mass which would be apt to injure the passengers or crew if such items became loose in the event of a minor crash landing, and the supporting structure shall be designed to restrain these items.

[15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1093, Feb. 5, 1952] § 4b.261 Structural ditching provisions. (For structural strength considerations of ditching provisions see § 4b.361 (c).)

[Amdt. 4b-8, 18 F. R. 2214, Apr. 18, 1953] FATIGUE EVALUATION General.

§ 4b.270

The strength, detail design, and fabrication of those portions of the airplane's flight structure in which fatigue may be critical shall be evaluated in accordance with the provisions of either paragraph (a) or (b) of this section.

(a) Fatigue strength. The structure shall be shown by analysis and/or tests to be capable of withstanding the repeated loads of variable magnitude expected in service. The provisions of subparagraphs (1) through (3) of this paragraph shall apply.

(1) Evaluation of fatigue shall involve the following:

(i) Typical loading spectrum expected in service;

(ii) Identification of principal structural elements and detail design points, the fatigue failure of which could cause catastrophic failure of the aircraft; and

(iii) An analysis and/or repeated load tests of principal structural elements and detail design points, identified in subdivision (ii) of this subparagraph;

NOTE: Usually tests of principal structural elements include major fittings, samples of Joints, spar cap strips, skin units, and other representative sections of the flight structure.

(2) It shall be acceptable to utilize the service history of airplanes of similar structural design, taking due account of

differences in operating conditions and procedures.

(3) When circumstances require substantiation of the pressure cabin by fatigue tests, the cabin or representative portions of it shall be cycle-pressure tested, utilizing the normal operating pressure together with the effects of external aerodynamic pressure combined with the flight loads. It shall be acceptable to represent the effects of flight loads by an increased cabin pressure, or to omit the flight loads if they are shown to have no significant effect upon fatigue.

(b) Fail safe strength. It shall be shown by analysis and/or tests that catastrophic failure or excessive structural deformation, which could adversely affect the flight characteristics of the airplane, are not probable after fatigue failure or obvious partial failure of a single principal structural element. After such failure, the remaining structure shall be capable of withstanding static loads corresponding with the flight loading condition specified in subparagraphs (1) through (4) of this paragraph. These loads shall be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise taken into consideration. In the case of a pressure cabin, the normal operating pressures combined with the expected external aerodynamic pressures shall be applied simultaneously with the flight loading conditions specified in this paragraph.

(1) An ultimate maneuver load factor of 2.0 at Vc.

(2) Gust loads as specified in §§ 4b.211 (b) and 4b.215 (b), except that these gust loads shall be considered to be ultimate and the gust velocities shall be as follows:

(i) At speed VB, 49 fps from sea level to 20,000 feet altitude, thereafter decreasing linearly to 28 fps at 50,000 feet altitude.

(ii) At speed Vc, 33 fps from sea level to 20,000 feet altitude, thereafter decreasing linearly to 16.5 fps at 50,000 > feet altitude.

(iii) At speed VD, 15 fps from sea level to 20,000 feet altitude, thereafter decreasing linearly to 6 fps at 50,000 feet altitude.

(3) Eighty percent of the limit loads resulting from the conditions specified in § 4b.220 (c). These loads shall be considered to be ultimate.

(4) Eighty percent of the limit maneuvering loads resulting from the conditions specified in § 4b.215 (a), except that the load need not exceed 100 percent of the critical load obtained in compliance with the provisions of § 4b.215 (a) using a pilot effort of 180 pounds. This load shall be considered to be ultimate. [Amdt. 4b-3, 21 F.R. 992, Feb. 11, 1956, as amended by Amdt. 4b-6, 22 F.R. 5564, July 16, 1957; Amdt. 4b-8, 23 F.R. 2590, Apr. 19, 1958]

§ 4b.270-1 Flight structure for fatigue evaluation (FAA interpretations which apply to § 4b.270).

The term "flight structure" as applied to fatigue evaluation is defined as those portions of the airplane's structure failure of which could result in catastrophic failure of the aircraft and includes the wings, fixed and movable control surfaces, fuselage, and their related primary attachments.

[Supp. 38, 23 F. R. 3031, May 7, 1958] § 4b.270-2 Fatigue evaluation, general (FAA policies which apply to § 4b.270).

The applicant should submit to the FAA a report outlining the procedures and the substantiating analyses and tests he proposes to follow in showing compliance with the fatigue evaluation requirements of § 4b.270. Typical procedures which may be used as guidance in the fatigue evaluation of the structures are discussed in Appendix H1 to the Civil Aeronautics Manual 4b. [Supp. 38, 23 F. R. 3031, May 7, 1958] Subpart D-Design and Construction

GENERAL

§ 4b.300 Scope.

1

use

on transport category airplanes. Where new or improved materials are used or where the materials are not covered by specifications sufficient information and data should be submitted to the Administrator to enable him to assess the suitability of the material. In all cases it is the responsibility of the applicant to demonstrate the adequacy of the materials employed.

[Supp. 25, 20 F. R. 2278, Apr. 8, 1955]

§ 4b.302 Fabrication methods.

The methods of fabrication employed in constructing the airplane structure shall be such as to produce a consistently sound structure. When a fabrication process such as gluing, spot welding, or heat treating requires close control to attain this objective, the process shall be performed in accordance with an approved process specification. § 4b.303

Standard fastenings.

salt water, particularly where parts made from different metals are in close proximity.

8 4b.305 Inspection provisions.

Means shall be provided to permit the close examination of those parts of the airplane which require periodic inspection, adjustment for proper alignment and functioning, and lubrication of moving parts.

§ 4b.306 Material strength properties and design values.

(a) Material strength properties shall be based on a sufficient number of tests of material conforming to specifications to establish design values on a statistical basis.

(b) The design values shall be so chosen that the probability of any structure being understrength because of material variations is extremely remote. The effects of temperature on allowable stresses used for design in an essential component or structure shall be considered where thermal effects are significant under normal operating conditions.

(c) ANC-5, ANC-18, and ANC-23 Part II values shall be used unless shown to be inapplicable in a particular case.

NOTE: ANC-5, "Strength of Metal Aircraft Elements," ANC-18, "Design of Wood Aircraft Structures," and ANC-23, "Sandwich Construction for Aircraft," are published by the Subcommittee on Air Force-Navy-Civil Aircraft Design Criteria, and may be obtained from the Superintendent of Documents, Government Printing Office, Washington 25, D. C.

(d) The strength, detail design, and fabrication of the structure shall be such as to minimize the probability of disastrous fatigue failure. (See also § 4b.270.)

NOTE: Points of stress concentration are one of the main sources of fatigue failure. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1093, Feb. 5, 1952; Amdt. 4b-8, 18 F. R. 2214, Apr. 18, 1953; Amdt. 4b-3, 21 F. R. 992, Feb. 11, 1956] § 4b.306-1

Material strength properties (FAA policies which apply to § 4b.306(c)).

(a) In the case of structures where the applied loads are eventually distributed through a single member within an assembly, the failure of which would result in the loss of the structural integrity of the component involved, the

guaranteed minimum design mechanical properties ("A" values) listed in ANC-5 should be used for design.

(b) Redundant structures wherein failure of individual elements would re|| sult in the applied load being safely distributed to other load carrying memIbers, may be designed on the basis of the "90 percent probability" ("B" values).

(c) When strength testing is employed to establish design allowables, such as in the case of sheet-stiffener compression tests, the test results should be reduced through use of a materials correction factor to values which would be met by material having the design allowable material properties for the part under consideration. The ANC-5 Bulletin outlines methods of accomplishing this reduction but these are by no means considered as the only methods available.

(d) Use of design values greater than the guaranteed minimums is permissible in applications where only guaranteed minimum values are normally permitted provided that the higher values are substantiated by "premium selection" of the material. These increased design allowables will be acceptable providing that a specimen or specimens of each individual item are tested prior to its use, to assure that the strength properties of the particular item will equal or exceed the properties to be used in design. Such quality control should also be exercised for the manufacture of spare parts.

[Supp. 25, 20 F. R. 2278, Apr. 8, 1955]

§ 4b.307 Special factors.

Where there is uncertainty concerning the actual strength of a particular part of the structure, or where the strength

a The ANC-5 Bulletin "Strength of Metal Aircraft Elements" specifies "A" and "B" values for allowable design properties. The "A" values are those which the material producer has indicated to be the minimum he expects for the given material. The only values considered guaranteed values are the tensile ultimate and tensile yield "A" values which have been published by the material producer for the grain direction accepted for commercial guarantees. The "B" values represent design properties which the materials producers have indicated will be met or exceeded by 90 percent of the material supplied by them. More detailed information on the derivation of related design mechanical properties can be obtained by referring to § 3.111 "Design Mechanical Properties" of the ANC-5 Bulletin.

is likely to deteriorate in service prior to normal replacement of the part, or where the strength is subject to appreciable variability due to uncertainties in manufacturing processes and inspection methods, the factor of safety prescribed in § 4b.200 (a) shall be multiplied by a special factor of a value such as to make the probability of the part being understrength from these causes extremely remote. The following special factors shall be used:

(a) Casting factors. (1) Where only visual inspection of a casting is to be employed, the casting factor shall be 2.0, except that it need not exceed 1.25 with respect to bearing stresses.

(2) It shall be acceptable to reduce the factor of 2.0 specified in subparagraph (1) of this paragraph to a value of 1.25 if such a reduction is substantiated by testing at least three sample castings and if the sample castings as well as all production castings are visually and radiographically inspected in accordance with an approved inspection specification. During these tests the samples shall withstand the ultimate load multiplied by the factor of 1.25 and in addition shall comply with the corresponding limit load multiplied by a factor of 1.15.

(3) Casting factors other than those contained in subparagraphs (1) and (2) of this paragraph shall be acceptable if they are found to be appropriately related to tests and to inspection proce dures.

(4) A casting factor need not be employed with respect to the bearing surface of a part if the bearing factor used (see paragraph (b) of this section) is of greater magnitude than the casting factor.

(b) Bearing factors. (1) Bearing factors shall be used of sufficient magnitude to provide for the effects of normal relative motion between parts and in joints with clearance (free fit) which are subject to pounding or vibration. (Bearing factor values for control surface and system joints are specified in §§ 4b.313 (a) and 4b.329 (b).)

(2) A bearing factor need not be employed on a part if another special factor prescribed in this section is of greater magnitude than the bearing factor.

(c) Fitting factors. (1) A fitting factor of at least 1.15 shall be used on all fittings the strength of which is not