260 

14 CFR Ch. I (1–1–24 Edition) 

§ 25.563 

would cause the greatest likelihood of 
the upper torso restraint system 
(where installed) moving off the occu-
pant’s shoulder, and with the wings 
level. Peak floor deceleration must 
occur in not more than 0.09 seconds 
after impact and must reach a min-
imum of 16g. Where floor rails or floor 
fittings are used to attach the seating 
devices to the test fixture, the rails or 
fittings must be misaligned with re-
spect to the adjacent set of rails or fit-
tings by at least 10 degrees vertically 
(

i.e., out of Parallel) with one rolled 10 

degrees. 

(c) The following performance meas-

ures must not be exceeded during the 
dynamic tests conducted in accordance 
with paragraph (b) of this section: 

(1) Where upper torso straps are used 

for crewmembers, tension loads in indi-
vidual straps must not exceed 1,750 
pounds. If dual straps are used for re-

straining the upper torso, the total 
strap tension loads must not exceed 
2,000 pounds. 

(2) The maximum compressive load 

measured between the pelvis and the 
lumbar column of the anthropomorphic 
dummy must not exceed 1,500 pounds. 

(3) The upper torso restraint straps 

(where installed) must remain on the 
occupant’s shoulder during the impact. 

(4) The lap safety belt must remain 

on the occupant’s pelvis during the im-
pact. 

(5) Each occupant must be protected 

from serious head injury under the con-
ditions prescribed in paragraph (b) of 
this section. Where head contact with 
seats or other structure can occur, pro-
tection must be provided so that the 
head impact does not exceed a Head In-
jury Criterion (HIC) of 1,000 units. The 
level of HIC is defined by the equation: 

HIC

t

t

t

t

a t dt

t

t

=

−

(

)

−

(

)

⎡

⎣

⎢

⎢

⎤

⎦

⎥

⎥

⎧

⎨

⎪

⎩⎪

⎫

⎬

⎪

⎭⎪

∫

2

1

2

1

2 5

1

1

2

( )

.

max

Where: 

t

1

is the initial integration time, 

t

2

is the final integration time, and 

a(t) is the total acceleration vs. time curve 

for the head strike, and where 

(t) is in seconds, and (a) is in units of gravity 

(g). 

(6) Where leg injuries may result 

from contact with seats or other struc-
ture, protection must be provided to 
prevent axially compressive loads ex-
ceeding 2,250 pounds in each femur. 

(7) The seat must remain attached at 

all points of attachment, although the 
structure may have yielded. 

(8) Seats must not yield under the 

tests specified in paragraphs (b)(1) and 
(b)(2) of this section to the extent they 
would impede rapid evacuation of the 
airplane occupants. 

[Amdt. 25–64, 53 FR 17646, May 17, 1988] 

§ 25.563

Structural ditching provi-

sions. 

Structural strength considerations of 

ditching provisions must be in accord-
ance with § 25.801(e). 

F

ATIGUE

E

VALUATION

 

§ 25.571

Damage-tolerance and fatigue 

evaluation of structure. 

(a) 

General.  An evaluation of the 

strength, detail design, and fabrication 
must show that catastrophic failure 
due to fatigue, corrosion, manufac-
turing defects, or accidental damage, 
will be avoided throughout the oper-
ational life of the airplane. This eval-
uation must be conducted in accord-
ance with the provisions of paragraphs 
(b) and (e) of this section, except as 
specified in paragraph (c) of this sec-
tion, for each part of the structure that 
could contribute to a catastrophic fail-
ure (such as wing, empennage, control 
surfaces and their systems, the fuse-
lage, engine mounting, landing gear, 
and their related primary attach-
ments). For turbojet powered air-
planes, those parts that could con-
tribute to a catastrophic failure must 
also be evaluated under paragraph (d) 
of this section. In addition, the fol-
lowing apply: 

VerDate Sep<11>2014 

09:06 Jun 28, 2024

Jkt 262046

PO 00000

Frm 00270

Fmt 8010

Sfmt 8010

Y:\SGML\262046.XXX

262046

EC28SE91.044</MATH>

jspears on DSK121TN23PROD with CFR

261 

Federal Aviation Administration, DOT 

§ 25.571 

(1) Each evaluation required by this 

section must include— 

(i) The typical loading spectra, tem-

peratures, and humidities expected in 
service; 

(ii) The identification of principal 

structural elements and detail design 
points, the failure of which could cause 
catastrophic failure of the airplane; 
and 

(iii) An analysis, supported by test 

evidence, of the principal structural 
elements and detail design points iden-
tified in paragraph (a)(1)(ii) of this sec-
tion. 

(2) The service history of airplanes of 

similar structural design, taking due 
account of differences in operating con-
ditions and procedures, may be used in 
the evaluations required by this sec-
tion. 

(3) Based on the evaluations required 

by this section, inspections or other 
procedures must be established, as nec-
essary, to prevent catastrophic failure, 
and must be included in the Airworthi-
ness Limitations section of the In-
structions for Continued Airworthiness 
required by § 25.1529. The limit of valid-
ity of the engineering data that sup-
ports the structural maintenance pro-
gram (hereafter referred to as LOV), 
stated as a number of total accumu-
lated flight cycles or flight hours or 
both, established by this section must 
also be included in the Airworthiness 
Limitations section of the Instructions 
for Continued Airworthiness required 
by § 25.1529. Inspection thresholds for 
the following types of structure must 
be established based on crack growth 
analyses and/or tests, assuming the 
structure contains an initial flaw of 
the maximum probable size that could 
exist as a result of manufacturing or 
service-induced damage: 

(i) Single load path structure, and 
(ii) Multiple load path ‘‘fail-safe’’ 

structure and crack arrest ‘‘fail-safe’’ 
structure, where it cannot be dem-
onstrated that load path failure, par-
tial failure, or crack arrest will be de-
tected and repaired during normal 
maintenance, inspection, or operation 
of an airplane prior to failure of the re-
maining structure. 

(b) 

Damage-tolerance evaluation. The 

evaluation must include a determina-
tion of the probable locations and 

modes of damage due to fatigue, corro-
sion, or accidental damage. Repeated 
load and static analyses supported by 
test evidence and (if available) service 
experience must also be incorporated 
in the evaluation. Special consider-
ation for widespread fatigue damage 
must be included where the design is 
such that this type of damage could 
occur. An LOV must be established 
that corresponds to the period of time, 
stated as a number of total accumu-
lated flight cycles or flight hours or 
both, during which it is demonstrated 
that widespread fatigue damage will 
not occur in the airplane structure. 
This demonstration must be by full- 
scale fatigue test evidence. The type 
certificate may be issued prior to com-
pletion of full-scale fatigue testing, 
provided the Administrator has ap-
proved a plan for completing the re-
quired tests. In that case, the Air-
worthiness Limitations section of the 
Instructions for Continued Airworthi-
ness required by § 25.1529 must specify 
that no airplane may be operated be-
yond a number of cycles equal to 

1

⁄

2

the 

number of cycles accumulated on the 
fatigue test article, until such testing 
is completed. The extent of damage for 
residual strength evaluation at any 
time within the operational life of the 
airplane must be consistent with the 
initial detectability and subsequent 
growth under repeated loads. The resid-
ual strength evaluation must show 
that the remaining structure is able to 
withstand loads (considered as static 
ultimate loads) corresponding to the 
following conditions: 

(1) The limit symmetrical maneu-

vering conditions specified in § 25.337 at 
all speeds up to V

c

and in § 25.345. 

(2) The limit gust conditions speci-

fied in § 25.341 at the specified speeds up 
to V

C

and in § 25.345. 

(3) The limit rolling conditions speci-

fied in § 25.349 and the limit unsymmet-
rical conditions specified in §§ 25.367 
and 25.427 (a) through (c), at speeds up 
to V

C

. 

(4) The limit yaw maneuvering condi-

tions specified in § 25.351(a) at the spec-
ified speeds up to V

C

. 

(5) For pressurized cabins, the fol-

lowing conditions: 

(i) The normal operating differential 

pressure combined with the expected 

VerDate Sep<11>2014 

09:06 Jun 28, 2024

Jkt 262046

PO 00000

Frm 00271

Fmt 8010

Sfmt 8010

Y:\SGML\262046.XXX

262046

jspears on DSK121TN23PROD with CFR

262 

14 CFR Ch. I (1–1–24 Edition) 

§ 25.581 

external aerodynamic pressures applied 
simultaneously with the flight loading 
conditions specified in paragraphs 
(b)(1) through (4) of this section, if they 
have a significant effect. 

(ii) The maximum value of normal 

operating differential pressure (includ-
ing the expected external aerodynamic 
pressures during 1 g level flight) multi-
plied by a factor of 1.15, omitting other 
loads. 

(6) For landing gear and directly-af-

fected airframe structure, the limit 
ground loading conditions specified in 
§§ 25.473, 25.491, and 25.493. 

If significant changes in structural 
stiffness or geometry, or both, follow 
from a structural failure, or partial 
failure, the effect on damage tolerance 
must be further investigated. 

(c) 

Fatigue (safe-life) evaluation. Com-

pliance with the damage-tolerance re-
quirements of paragraph (b) of this sec-
tion is not required if the applicant es-
tablishes that their application for par-
ticular structure is impractical. This 
structure must be shown by analysis, 
supported by test evidence, to be able 
to withstand the repeated loads of vari-
able magnitude expected during its 
service life without detectable cracks. 
Appropriate safe-life scatter factors 
must be applied. 

(d) 

Sonic fatigue strength. It must be 

shown by analysis, supported by test 
evidence, or by the service history of 
airplanes of similar structural design 
and sonic excitation environment, 
that— 

(1) Sonic fatigue cracks are not prob-

able in any part of the flight structure 
subject to sonic excitation; or 

(2) Catastrophic failure caused by 

sonic cracks is not probable assuming 
that the loads prescribed in paragraph 
(b) of this section are applied to all 
areas affected by those cracks. 

(e) 

Damage-tolerance (discrete source) 

evaluation.  The airplane must be capa-
ble of successfully completing a flight 
during which likely structural damage 
occurs as a result of— 

(1) Impact with a 4-pound bird when 

the velocity of the airplane relative to 
the bird along the airplane’s flight 
path is equal to V

c

at sea level or 0.85V

c

 

at 8,000 feet, whichever is more critical; 

(2) Uncontained fan blade impact; 
(3) Uncontained engine failure; or 

(4) Uncontained high energy rotating 

machinery failure. 

The damaged structure must be able to 
withstand the static loads (considered 
as ultimate loads) which are reason-
ably expected to occur on the flight. 
Dynamic effects on these static loads 
need not be considered. Corrective ac-
tion to be taken by the pilot following 
the incident, such as limiting maneu-
vers, avoiding turbulence, and reducing 
speed, must be considered. If signifi-
cant changes in structural stiffness or 
geometry, or both, follow from a struc-
tural failure or partial failure, the ef-
fect on damage tolerance must be fur-
ther investigated. 

[Amdt. 25–45, 43 FR 46242, Oct. 5, 1978, as 
amended by Amdt. 25–54, 45 FR 60173, Sept. 
11, 1980; Amdt. 25–72, 55 FR 29776, July 20, 
1990; Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; 
Amdt. 25–96, 63 FR 15714, Mar. 31, 1998; 63 FR 
23338, Apr. 28, 1998; Amdt. 25–132, 75 FR 69781, 
Nov. 15, 2010; Amdt. No. 25–148, 87 FR 75710, 
Dec. 9, 2022; 88 FR 2813, Jan. 18, 2023] 

L

IGHTNING

P

ROTECTION

 

§ 25.581

Lightning protection. 

(a) The airplane must be protected 

against catastrophic effects from light-
ning. 

(b) For metallic components, compli-

ance with paragraph (a) of this section 
may be shown by— 

(1) Bonding the components properly 

to the airframe; or 

(2) Designing the components so that 

a strike will not endanger the airplane. 

(c) For nonmetallic components, 

compliance with paragraph (a) of this 
section may be shown by— 

(1) Designing the components to min-

imize the effect of a strike; or 

(2) Incorporating acceptable means of 

diverting the resulting electrical cur-
rent so as not to endanger the airplane. 

[Amdt. 25–23, 35 FR 5674, Apr. 8, 1970] 

Subpart D—Design and 

Construction 

G

ENERAL

 

§ 25.601

General. 

The airplane may not have design 

features or details that experience has 
shown to be hazardous or unreliable. 
The suitability of each questionable 

VerDate Sep<11>2014 

09:06 Jun 28, 2024

Jkt 262046

PO 00000

Frm 00272

Fmt 8010

Sfmt 8010

Y:\SGML\262046.XXX

262046

jspears on DSK121TN23PROD with CFR