14 CFR Part 25, Subpart C - Structure

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. Typical Interpretation: For cabin interiors sizing is almost always done to ultimate loads (crash or emergency landing, refer to 14 CFR Part 25.561). There are situations where in flight load factors may be used, but even these can be ultimate 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 airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. 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. Typical Interpretation: Inertia loads must be applied on structures. Loads and moments must be in equilibrium with reactions at the aircraft structure interface points. Acceptable load application methods (simulated and real world) are body loads, pressure loads (using load spreaders) and point loads where applicable. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account. Typical Interpretation: Large deflections must be accounted for using nonlinear geometry effects and differential stiffness effects (FEM). This is rare in cabin interiors, but if needed, it will help reduce total deflection and also account for any load redistribution at the aircraft interface points due to large deflections. 25.303 Factor of Safety Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit load which is considered external loads on the structure. When a loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified. Typical Interpretation: Except in rare cases where limit loads are considered, all loads in cabin interiors are ultimate loads, including abuse loads (not required by FAA). No 1.5 factor is required on the applied loads.

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. Typical Interpretation: Except in rare cases where limit loads are considered, all loads in cabin interiors are ultimate loads including abuse loads (not required by FAA). The emphasis in cabin interiors is on safe egress of passengers or no injury in case of inflight load cases. Permanent deflection in ultimate cases is allowed as long as it is not detrimental to safe passenger egress. Usually, the margins of safety are written against ultimate allowable values. If the ratio of the ultimate allowable to yield allowable is less than 1.5, no permanent deformation need be expected for limit loads (ultimate load/1.5). (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.

Typical Interpretation: In full scale structure static testing (you will learn about this in advanced level courses), the total load is applied using what is known as a “whiffle tree”. This max ultimate load is held for 3 seconds and deflections are measured at critical locations, usually at the corners and mid height locations. When the certification is done by analysis, then the modeling techniques and the modeling process must have been through an FEM validation process. For example, the customer or client FEM guidelines specification document (or structural design criteria document etc.) may require that the material properties should be such that the actual sandwich panel coupon test deflection measurements match the simulation deflections of the same coupon. Similarly, every client company may have their own requirements in their SDC ultimately to satisfy the regulation above. (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. Typical Interpretation: This is only applicable to resonance or amplification due to vibration. Cabin interiors are certified to ultimate loads which are essentially quasi static equivalent load factors due to a dynamic event such as an emergency landing or crash or rapid decompression events. However, there may be other dynamic load cases such as sustained engine imbalance (SEI) that require additional analysis. This is generally not required for a majority of installations, at least in my experience. This topic is currently out of the scope of Stress Ebook LLC. However, I might get into it in the future and create a special course if there is enough demand for it. (d) [Reserved] (e) The airplane 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 Administrator. Typical Interpretation:

For cabin interiors, these load factors may be interpreted as ultimate flight load factors. Generally these factors will be prescribed by the aircraft OEM or the aircraft integrator/supplier SDC. (f) Unless shown to be extremely improbable, the airplane must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These must be considered limit loads and must be investigated at airspeeds up to VC/MC. Typical Interpretation: For cabin interiors, these load factors may be interpreted as ultimate flight load factors. Generally these factors will be prescribed by the aircraft OEM or the aircraft integrator/supplier SDC.

25.307 Proof of Strength (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. The Administrator may require ultimate load tests in cases where limit load tests may be inadequate. Typical Interpretation: Compliance is shown by static testing or analysis based on validated FEM modeling techniques and industry standard classical hand calculations. (b) -(c) [Reserved] (d) When static or dynamic tests are used to show compliance with the requirements of § 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. Typical Interpretation: Compliance to this regulation is shown by considering various factors for static testing, such as load factor of 1.15 on total applied load, load factor of 1.33 for items subject to wear and tear (latches, quarter turns), load factor of 1.33 for seat track attachment hardware etc., this is also linked to 25.613 thru 25.625 below. For certification by analysis, fitting and special factors are used, compliance is demonstrated using A and B basis allowable loads and stresses. A-Basis (more stringent statistically and lower values) is used for sizing the components of a load path that has no redundancy, in other words multiple load paths in case of failure of one link. B-Basis values can be used for components of a load path that has redundant load path features, example panel pins or joints with multiple inserts in the joint.

25.561 Emergency Landing Conditions (a) The airplane, although it may be damaged in emergency landing conditions on land or water, must be designed as prescribed in this section to protect each occupant under those conditions. Typical Interpretation: All cabin structures must be designed and certified in such a manner so as to not impede safe egress of the passengers in case of emergency landing or crash landing. (b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury in a minor crash landing when—

(1) Proper use is made of seats, belts, and all other safety design provisions; (2) The wheels are retracted (where applicable); and (3) The occupant experiences the following ultimate inertia forces acting separately relative to the surrounding structure:

(i) Upward, 3.0g (ii) Forward, 9.0g (iii) Sideward, 3.0g on the airframe; and 4.0g on the seats and their attachments. (iv) Downward, 6.0g (v) Rearward, 1.5g

Typical Interpretation: All cabin structures must be certified, at a minimum, to the above inertia load cases. In addition, flight load cases, combination load cases, rapid decompression load cases and abuse load cases may also be required. Some exceptions to this rule maybe certain structures that do not lie along the safe egress path, in certain directions. Example may be a bulkhead in the nose area. The FWD load case may not be required for this bulkhead. However a lot of times rapid decompression cases are more critical. (c) For equipment, cargo in the passenger compartments and any other large masses, the following apply:

(1) Except as provided in paragraph (c)(2) of this section, these items must be positioned so that if they break loose they will be unlikely to:

(i) Cause direct injury to occupants; (ii) Penetrate fuel tanks or lines or cause fire or explosion hazard by damage to adjacent systems; or (iii) Nullify any of the escape facilities provided for use after an emergency landing.

Typical Interpretation: Every insert, loose contents or items of mass in the cabin interior structures must be restrained using restraint devices such as structure compartment panels, dual door latches, quarter turns etc. Certain restraint devices must be certified to 1.33 times the applied load due to frequent use or wear and tear, such as quarter turns.

(2) When such positioning is not practical (e.g. fuselage mounted engines or auxiliary power units) each such item of mass shall be restrained under all loads up to those specified in paragraph (b)(3) of this section. The local attachments for these items should be designed to withstand 1.33 times the specified loads if these items are subject to severe wear and tear through frequent removal (e.g. quick change interior items).

Typical Interpretation: Every insert, loose contents or items of mass in the cabin interior structures must be restrained using restraint devices such as structure compartment panels, dual door latches, quarter turns etc. Certain restraint devices must be certified to 1.33 times the applied load due to frequent use or wear and tear, such as quarter turns.

(d) Seats and items of mass (and their supporting structure) must not deform under any loads up to those specified in paragraph (b)(3) of this section in any manner that would impede subsequent rapid evacuation of occupants. Typical Interpretation: Every insert, loose contents or items of mass in the cabin interior structures must be restrained using restraint devices such as structure compartment panels, dual door latches, quarter turns etc. Certain restraint devices must be certified to 1.33 times the applied load due to frequent use or wear and tear, such as quarter turns. 25.563 Structural Ditching Provisions Structural strength considerations of ditching provisions must be in accordance with § 25.801(e). § 25.801 (a) If certification with ditching provisions is requested, the airplane must meet the requirements of this section and §§ 25.807(e), 25.1411, and 25.1415(a). (b) Each practicable design measure, compatible with the general characteristics of the airplane, must be taken to minimize the probability that in an emergency landing on water, the behavior of the airplane would cause immediate injury to the occupants or would make it impossible for them to escape. (c) The probable behavior of the airplane in a water landing must be investigated by model tests or by comparison with airplanes of similar configuration for which the ditching characteristics are known. Scoops, flaps, projections, and any other factor likely to affect the hydrodynamic characteristics of the airplane, must be considered. (d) It must be shown that, under reasonably probable water conditions, the flotation time and trim of the airplane will allow the occupants to leave the airplane and enter the liferafts required by § 25.1415. If compliance with this provision is shown by buoyancy and trim computations, appropriate allowances must be made for probable structural damage and leakage. If the airplane has fuel tanks (with fuel jettisoning provisions) that can reasonably be expected to withstand a ditching without leakage, the jettisonable volume of fuel may be considered as buoyancy volume. (e) Unless the effects of the collapse of external doors and windows are accounted for in the investigation of the probable behavior of the airplane in a water landing (as prescribed in paragraphs (c) and (d) of this section), the external doors and windows must be designed to withstand the probable maximum local pressures. Typical Interpretation: Ditching provisions are usually included for planes that are flying over waters and certification is for ditching is required for such aircraft. Usually these provisions are in the form of handles at the structures next to the FWD and AFT (and sometimes middle) door locations. These provisions must be certified to ditching loads from a typical life raft attached to them via the life raft cable. The load can be 1000lb or more.

14 CFR Part 25, Subpart D – Design and Construction 25.603 Materials The suitability and durability of materials used for parts, the failure of which could adversely affect safety, must— (a) Be established on the basis of experience or tests; (b) Conform to approved specifications (such as industry or military specifications, or Technical Standard Orders) that ensure their having the strength and other properties assumed in the design data; and (c) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service. Typical Interpretation: Materials are typically restricted based on procurement specifications. Usually NAS MS specification parts and parts noted in the MMPDS document are used. Panels made according to FAA approved processes and materials are used. Exceptions may be used upon approval of the proper certifying authorities (FAA, DER, ODA Unit Members etc.).

25.605 Fabrication Methods (a) The methods of fabrication used must produce a consistently sound structure. If a fabrication process (such as gluing, spot welding, or heat treating) requires close control to reach this objective, the process must be performed under an approved process specification. (b) Each new aircraft fabrication method must be substantiated by a test program. Typical Interpretation: Fitting factors take care of any uncertainty in the metal parts material variation. Other fabrication or manufacturing processes for bonded structures and panels are a major and critical part of any cabin interior manufacturer. These specifications must have been FAA approved. Panels fabricated according to FAA approved processes do not need an uncertainty factor. Similarly assemblies (example standard tie rod assemblies) with tested allowable loads may not need a factor. Exceptions may be used upon approval of the proper certifying authorities (FAA, DER, ODA Unit Members etc.). 25.611 Accessibility Provisions (a) Means must be provided to allow inspection (including inspection of principal structural elements and control systems), replacement of parts normally requiring replacement, adjustment, and lubrication as necessary for continued airworthiness. The inspection means for each item must be practicable for the inspection interval for the item. Nondestructive inspection aids may be used to inspect structural elements where it is impracticable to provide means for direct visual inspection if it is shown that the inspection is effective and the inspection procedures are specified in the maintenance manual required by § 25.1529. (b) EWIS must meet the accessibility requirements of § 25.1719 Typical Interpretation: Typically these provisions are included in the form of panel cutouts and cover plates, which must be accounted for in the FEM modeling process.

25.619 Special Factors The factor of safety prescribed in § 25.303 must be multiplied by the highest pertinent special factor of safety prescribed in §§ 25.621 through 25.625 for each part of the structure whose strength is—

(a) Uncertain; (b) Likely to deteriorate in service before normal replacement; or (c) Subject to appreciable variability because of uncertainties in manufacturing processes or inspection methods.

Typical Interpretation: Fitting factor (1.15) is required for aircraft primary structure installation hardware margins. Parts

with tested allowable loads moments or stresses may not need this factor, but it depends on the

approving authority (DER, UM, ODA etc.) also. Panels typically do not need this factor if they are

manufactured according the manufacturer’s FAA certified process specifications.

25.621 Casting Factors (a) General. The factors, tests, and inspections specified in paragraphs (b) through (d) of this section must be applied in addition to those necessary to establish foundry quality control. The inspections must meet approved specifications. Paragraphs (c) and (d) of this section apply to any structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads. (b) Bearing stresses and surfaces. The casting factors specified in paragraphs (c) and (d) of this section—

(1) Need not exceed 1.25 with respect to bearing stresses regardless of the method of inspection used; and (2) Need not be used with respect to the bearing surfaces of a part whose bearing factor is larger than the applicable casting factor.

(c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the airplane or result in serious injury to occupants, the following apply:

(1) Each critical casting must— (i) Have a casting factor of not less than 1.25; and (ii) Receive 100 percent inspection by visual, radiographic, and magnetic particle or penetrant inspection methods or approved equivalent nondestructive inspection methods.

(2) For each critical casting with a casting factor less than 1.50, three sample castings must be static tested and shown to meet—

(i) The strength requirements of § 25.305 at an ultimate load corresponding to a casting factor of 1.25; and (ii) The deformation requirements of § 25.305 at a load of 1.15 times the limit load.

(3) Examples of these castings are structural attachment fittings, parts of flight control systems, control surface hinges and balance weight attachments, seat, berth, safety belt, and fuel and oil tank supports and attachments, and cabin pressure valves.

(d) Noncritical castings. For each casting other than those specified in paragraph (c) of this section, the following apply:

(1) Except as provided in paragraphs (d)(2) and (3) of this section, the casting factors and corresponding inspections must meet the following table: Casting factor Inspection 2.0 or more 100 percent visual. Less than 2.0 but more than 1.5 100 percent visual, and magnetic particle or penetrant or equivalent nondestructive inspection methods. 1.25 through 1.50 100 percent visual, magnetic particle or penetrant, and radiographic, or approved equivalent nondestructive inspection methods.

(2) The percentage of castings inspected by nonvisual methods may be reduced below that specified in paragraph (d)(1) of this section when an approved quality control procedure is established. (3) For castings procured to a specification that guarantees the mechanical properties of the material in the casting and provides for demonstration of these properties by test of coupons cut from the castings on a sampling basis—

(i) A casting factor of 1.0 may be used; and (ii) The castings must be inspected as provided in paragraph (d)(1) of this section for casting factors of “1.25 through 1.50” and tested under paragraph (c)(2) of this section.

Typical Interpretation: Fitting factor (1.15) is required for aircraft primary structure installation hardware margins. Seat

track fittings require a 1.33 factor on loads, parts with tested allowable loads moments or stresses

may not need this factor, but it depends on the approving authority (DER, UM, ODA etc.) also.

Panels typically do not need this factor if they are manufactured according the manufacturer’s

FAA certified process specifications. A 1.33 factor is required for restraint devices experiencing

frequent operation such as quarter turns. Cargo net attachment provisions also require a 1.33

factor.

25.623 Bearing Factors

(a) Except as provided in paragraph (b) of this section, each part that has clearance (free fit), and that is subject to pounding or vibration, must have a bearing factor large enough to provide for the effects of normal relative motion. (b) No bearing factor need be used for a part for which any larger special factor is prescribed. Typical Interpretation: This is almost always avoided using properly sized bushings to minimize or eliminate any clearances. Typical examples would be tie rod end to lug joints. 25.625 Fitting Factors For each fitting (a part or terminal used to join one structural member to another), the following apply: (a) For each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1.15 must be applied to each part of—

(1) The fitting; (2) The means of attachment; and (3) The bearing on the joined members.

(b) No fitting factor need be used— (1) For joints made under approved practices and based on comprehensive test data (such as continuous joints in metal plating, welded joints, and scarf joints in wood); or (2) With respect to any bearing surface for which a larger special factor is used.

(c) For each integral fitting, the part must be treated as a fitting up to the point at which the section properties become typical of the member. (d) For each seat, berth, safety belt, and harness, the fitting factor specified in § 25.785(f)(3) applies. Typical Interpretation: Fitting factor (1.15) is required for aircraft primary structure installation hardware margins on the

joint components such as the fitting bolt, fitting, washers, screws and inserts in the panel.

25.787 Stowage Compartments (a) Each compartment for the stowage of cargo, baggage, carry-on articles, and equipment (such as life rafts), and any other stowage compartment must be designed for its placarded maximum weight of contents and for the critical load distribution at the appropriate maximum load factors corresponding to the specified flight and ground load conditions, and to the emergency landing conditions of § 25.561(b), except that the forces specified in the emergency landing conditions need not be applied to compartments located below, or forward, of all occupants in the airplane. If the airplane has a passenger seating configuration, excluding pilots seats, of 10 seats or more, each stowage compartment in the passenger cabin, except for underseat and overhead compartments for passenger convenience, must be completely enclosed. (b) There must be a means to prevent the contents in the compartments from becoming a hazard by shifting, under the loads specified in paragraph (a) of this section. For stowage compartments in the passenger and crew cabin, if the means used is a latched door, the design must take into consideration the wear and deterioration expected in service. (c) If cargo compartment lamps are installed, each lamp must be installed so as to prevent contact between lamp bulb and cargo. Typical Interpretation: All restraint devices including structure panels, doors, door latches, quarter turns, attachment inserts, installation hardware etc. must be certified to the loads and factors discussed before. 25.789 Retention of items of mass in passenger and crew compartments and galleys (a) Means must be provided to prevent each item of mass (that is part of the airplane type design) in a passenger or crew compartment or galley from becoming a hazard by shifting under the appropriate maximum load factors corresponding to the specified flight and ground load conditions, and to the emergency landing conditions of § 25.561(b). (b) Each interphone restraint system must be designed so that when subjected to the load factors specified in § 25.561(b)(3), the interphone will remain in its stowed position. Typical Interpretation: All restraint devices including structure panels, doors, door latches, quarter turns, attachment inserts, installation hardware etc. must be certified to the loads and factors discussed before.