FAR Structural Design Criteria
Subpart C--Structure
General
Sec. 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.
(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.
(c) If deflections under load would significantly change the distribution
of external or internal loads, this redistribution must be taken into
account.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.303 Factor of safety.
Unless otherwise specified, a factor of safety of 1.5 must be applied to
the prescribed limit load which are 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.
[Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]
Sec. 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.
(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.
(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.
(d) The dynamic response of the airplane to vertical and lateral continuous
turbulence must be taken into account. The continuous gust design criteria of
Appendix G of this part must be used to establish the dynamic response unless
more rational criteria are shown.
(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.
(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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-77, 57
FR 28949, June 29, 1992]
*****************************************************************************
57 FR 28946, No. 125, June 29, 1992
SUMMARY: This amendment revises the airworthiness standards of the Federal
Aviation Regulations (FAR) for type certification of transport category
airplanes concerning vibration, buffet, flutter and divergence. It clarifies
the requirement to consider flutter and divergence when treating certain
damage and failure conditions required by other sections of the FAR and
adjusts the safety margins related to aeroelastic stabiity to make them more
appropriate for the conditions to which they apply. These changes are made to
provide consistency with other sections of the FAR and to take into account
advances in technology and the evolution of the design of transport
airplanes.
EFFECTIVE DATE: July 29, 1992.
*****************************************************************************
Sec. 25.307 Proof of structure.
(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.
(b) [Reserved]
(c) [Reserved]
(d) When static or dynamic tests are used to show compliance with the
requirements of Sec. 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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55
FR 29775, July 20, 1990]
Flight Loads
Sec. 25.321 General.
(a) Flight load factors represent the ratio of the aerodynamic force
component (acting normal to the assumed longitudinal axis of the airplane) to
the weight of the airplane. A positive load factor is one in which the
aerodynamic force acts upward with respect to the airplane.
(b) Considering compressibility effects at each speed, compliance with the
flight load requirements of this subpart must be shown--
(1) At each critical altitude within the range of altitudes selected by the
applicant;
(2) At each weight from the design minimum weight to the design maximum
weight appropriate to each particular flight load condition; and
(3) For each required altitude and weight, for any practicable distribution
of disposable load within the operating limitations recorded in the Airplane
Flight Manual.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Flight Maneuver and Gust Conditions
Sec. 25.331 General.
(a) Procedure. The analysis of symmetrical flight must include at least the
conditions specified in paragraphs (b) through (d) of this section. The
following procedure must be used:
(1) Enough points on the maneuvering and gust envelopes must be
investigated to ensure that the maximum load for each part of the airplane
structure is obtained. A conservative combined envelope may be used.
(2) The significant forces acting on the airplane must be placed in
equilibrium in a rational or conservative manner. The linear inertia forces
must be considered in equilibrium with thrust and all aerodynamic loads,
while the angular (pitching) inertia forces must be considered in equilibrium
with thrust and all aerodynamic moments, including moments due to loads on
components such as tail surfaces and nacelles. Critical thrust values in the
range from zero to maximum continuous thrust must be considered.
(3) Where sudden displacement of a control is specified, the assumed rate
of control surface displacement may not be less than the rate that could be
applied by the pilot through the control system.
(4) In determining elevator angles and chordwise load distribution (in the
maneuvering conditions of paragraphs (b) and (c) of this section) in turns
and pull-ups, the effect of corresponding pitching velocities must be taken
into account. The in-trim and out-of-trim flight conditions specified in Sec.
25.255 must be considered.
(b) Maneuvering balanced conditions. Assuming the airplane to be in
equilibrium with zero pitching acceleration, the maneuvering conditions A
through I on the maneuvering envelope in Sec. 25.333(b) must be investigated.
(c) Maneuvering pitching conditions. The following conditions involving
pitching acceleration must be investigated:
(1) Maximum elevator displacement at VA. The airplane is assumed to be
flying in steady level flight (point A1, Sec. 25.333(b)) and, except as
limited by pilot effort in accordance with Sec. 25.397(b), the pitching
control is suddenly moved to obtain extreme positive pitching acceleration
(nose up). The dynamic response or, at the option of the applicant, the
transient rigid body response of the airplane, must be taken into account in
determining the tail load. Airplane loads which occur subsequent to the
normal acceleration at the center of gravity exceeding the maximum positive
limit maneuvering load factor, n, need not be considered.
(2) Specified control displacement. A checked maneuver, based on a rational
pitching control motion vs. time profile, must be established in which the
design limit load factor specified in Sec. 25.337 will not be exceeded.
Unless lesser values cannot be exceeded, the airplane response must result in
pitching accelerations not less than the following:
(i) A positive pitching acceleration (nose up) is assumed to be reached
concurrently with the airplane load factor of 1.0 (Points A1 to D1, Sec.
25.333(b)). The positive acceleration must be equal to at least
39n
---- (n-1.5), (Radians/sec./2/ )
v
where--
n is the positive load factor at the speed under consideration, and V is the
airplane equivalent speed in knots.
(ii) A negative pitching acceleration (nose down) is assumed to be reached
concurrently with the positive maneuvering load factor (Points A2 to D2,
Sec. 25.333(b)). This negative pitching acceleration must be equal to at
least
-26n
------ (n-1.5), (Radians/sec./2/ )
v
where--
n is the positive load factor at the speed under consideration; and V is the
airplane equivalent speed in knots.
(d) Gust conditions. The gust conditions B' through J' Sec. 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 Kg may be applied to the specified gust intensity for the
horizontal tail.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50594, Oct. 30, 1978; 43 FR 52495,
Nov. 13, 1978; 43 FR 54082, Nov. 20, 1978; Amdt. 25-72, 55 FR 29775, July 20,
1990; 55 FR 37607, Sept. 12, 1990]
Sec. 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 Sec. 25.1501.
(b) Maneuvering envelope.
[ ...Illustration appears here... ]
(c) Gust envelope.
[ ...Illustration appears here... ]
Sec. 25.335 Design airspeeds.
The selected design airspeeds are equivalent airspeeds (EAS). Estimated
values of VS0 and VS1 must be conservative.
(a) Design cruising speed, VC. For VC, the following apply:
(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, VC may be limited to a
selected Mach number.
(b) Design dive speed, VD. VD must be selected so that VC/MC is not greater
than 0.8 VD/MD, or so that the minimum speed margin between VC/MC and VD/MD
is the greater of the following values:
(1) From an initial condition of stabilized flight at VC/MC, the airplane
is upset, flown for 20 seconds along a flight path 7.5 deg. below the initial
path, and then pulled up at a load factor of 1.5 g (0.5 g acceleration
increment). The speed increase occurring in this maneuver may be calculated
if reliable or conservative aerodynamic data is used. Power as specified in
Sec. 25.175(b)(1)(iv) is assumed until the pull-up is initiated, at which time
power reduction and the use of pilot controlled drag devices may be assumed;
(2) The minimum speed margin must be enough to provide for atmospheric
variations (such as horizontal gusts, and penetration of jet streams and cold
fronts) and for instrument errors and airframe production variations. These
factors may be considered on a probability basis. However, the margin at
altitude where MC is limited by compressibility effects may not be less than
0.05 M.
(c) Design maneuvering speed VA. For VA, the following apply:
(1) VA may not be less than VS1 <radical>n where--
(i) n is the limit positive maneuvering load factor at VC; and
(ii) VS1 is the stalling speed with flaps retracted.
(2) VA and VS must be evaluated at the design weight and altitude under
consideration.
(3) VA need not be more than VC or the speed at which the positive CN max
curve intersects the positive maneuver load factor line, whichever is less.
(d) Design speed for maximum gust intensity, VB. For VB, the following
apply:
(1) VB may not be less than the speed determined by the intersection of the
line representing the maximum position lift CN max and the line representing
the rough air gust velocity on the gust V-n diagram, or (<radical>ng) VS1,
whichever is less, where--
(i) ng is the positive airplane gust load factor due to gust, at speed VC
(in accordance with Sec. 25.341), and at the particular weight under
consideration; and
(ii) VS1 is the stalling speed with the flaps retracted at the particular
weight under consideration.
(2) VB 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 Sec. 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) VF may not be less than--
(i) 1.6 VS1 with the flaps in takeoff position at maximum takeoff weight;
(ii) 1.8 VS1 with the flaps in approach position at maximum landing weight,
and
(iii) 1.8 VS0 with the flaps in landing position at maximum landing weight.
(f) Design drag device speeds, VDD. The selected design speed for each drag
device must be sufficiently greater than the speed recommended for the
operation of the device to allow for probable variations in speed control.
For drag devices intended for use in high speed descents, VDD may not be less
than VD. When an automatic drag device positioning or load limiting means is
used, the speeds and corresponding drag device positions programmed or
allowed by the automatic means must be used for design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.337 Limit maneuvering load factors.
(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 Vn
may not be less than 2.1+24,000/ (W +10,000) except that "n" may not be less
than 2.5 and need not be greater than 3.8--where "W" is the design maximum
takeoff weight.
(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 VC 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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.341 Gust loads.
(a) 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:
(1) Positive (up) and negative (down) rough air gusts of 66 fps at VB must
be considered at altitudes between sea level and 20,000 feet. The gust
velocity may be reduced linearly from 66 fps at 20,000 feet to 38 fps at
50,000 feet.
(2) Positive and negative gusts of 50 fps at VC must be considered at
altitudes between sea level and 20,000 feet. The gust velocity may be reduced
linearly from 50 fps at 20,000 feet to 25 fps at 50,000 feet.
(3) Positive and negative gusts of 25 fps at VD must be considered at
altitudes between sea level and 20,000 feet. The gust velocity may be reduced
linearly from 25 fps at 20,000 feet to 12.5 fps at 50,000 feet.
(b) The following assumptions must be made:
(1) The shape of the gust is
Ude 2<pi>s
U = --- (1-cos ------ )
2 25C
where--
s=distance penetrated into gust (ft);
C=mean geometric chord of wing (ft); and
Ude=derived gust velocity referred to in paragraph (a) (fps).
(2) Gust load factors vary linearly between the specified conditions B'
through G', as shown on the gust envelope in Sec. 25.333(c).
(c) In the absence of a more rational analysis, the gust load factors must
be computed as follows:
KgUdeVa
n=1 + ---------
498 (W/S)
where--
0.88<mu>g
Kg = ----------- = gust alleviation factor;
5.3+<mu>g
2(W/S)
<mu>g = ------ = airplane mass ratio:
rCag
Ude=derived gust velocities referred to in paragraph (a) (fps);
r=density of air (slugs cu. ft.);
W/S=wing 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 method. The wing lift
curve slope CAL 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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990]
Sec. 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 Sec. 25.1001(e) and (f), as applicable, 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 Sec.
25.341; and
(2) Fatigue evaluation of the structure must account for any increase in
operating stresses resulting from the design condition of paragraph (b)(1) of
this section; and
(3) The flutter, deformation, and vibration requirements must also be met
with zero fuel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 33 FR
12226, Aug. 30, 1968; Amdt. 25-72, 55 FR 29775, July 20, 1990; 55 FR 37607,
Sept. 12, 1990]
Sec. 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 Sec.
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 VF, 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 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 is assumed to be
subjected to symmetrical maneuvers and gusts within the range determined by--
(1) Maneuvering to a positive limit load factor as prescribed in Sec.
25.337(b); and
(2) Positive and negative derived gusts as prescribed in Sec. 25.341 acting
normal to the flight path in level flight.
(d) The airplane must be designed for landing at the maximum takeoff weight
with a maneuvering load factor of 1.5g and the flaps and similar high lift
devices in the landing configuration.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 29775, July 20, 1990; 55 FR 37607,
Sept. 12, 1990]
Sec. 25.349 Rolling 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 Sec. 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 maneuver.
(2) At VA, a sudden deflection of the aileron to the stop is assumed.
(3) At VC, the aileron deflection must be that required to produce a rate
of roll not less than that obtained in paragraph (a)(2) of this section.
(4) At VD, the aileron deflection must be that required to produce a rate
of roll not less than one-third of that in paragraph (a)(2) of this section.
(b) Unsymmetrical gusts. The condition of unsymmetrical gusts must be
considered by modifying the symmetrical flight conditions B' or C' (in Sec.
25.333(c)) whichever produces the critical load. 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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 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 VMC to VD, 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 surface stops, or by a 300-pound rudder pedal force,
whichever is less.
(2) With the rudder deflected as specified in paragraph (a)(1) of this
section, 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 paragraph (a)(1) of this section, 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 Sec.
25.333(c)) (as determined by Secs. 25.341 and 25.345(a)(2) or Sec.
25.345(c)(2)) must be investigated. The shape of the gust must be as
specified in Sec. 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:
KgtUdeVatSt
Lt = -----------
498
where--
Lt=vertical tail load (lbs.);
0.88<mu>gt
Kgt = ------------ = gust alleviation factor;
5.3+<mu>gt
2W K
<mu>gt = -------- ( ----- )**2 =lateral mass ratio;
pCtgatSt lt
Ude=derived gust velocity (fps);
p=air density (slugs/cu. ft.);
W=airplane weight (lbs.);
St=area of vertical tail (ft.**2);
Ct=mean geometric chord of vertical surface (ft.);
at=lift curve slope of vertical tail (per radian);
K=radius of gyration in yaw (ft).;
lt=distance from airplane c.g., to lift center of vertical surface (ft.);
g=acceleration due to gravity (ft./sec.**2); and
V=airplane equivalent speed (knots).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55
FR 29775, July 20, 1990; 55 FR 37608, Sept. 12, 1990; 55 FR 41415, Oct. 11,
1990]