ASTM F3116/F3116M-23a

This international standard was developed in accordance with internationally recognized principles on standardization established in the Decision on Principles for the
Development of International Standards, Guides and Recommendations issued by the World Trade Organization Technical Barriers to Trade (TBT) Committee.
Designation: F3116/F3116M − 23a
Standard Specification for
Design Loads and Conditions
This standard is issued under the fixed designation F3116/F3116M; the number immediately following the designation indicates the year
of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval.
A superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope 2. Referenced Documents
1.1 This specification addresses the airworthiness require- 2.1 ASTM Standards:
F3060 Terminology for Aircraft
ments for the design loads and conditions of small airplanes.
F3331 Practice for Aircraft Water Loads
1.2 This specification is applicable to small airplanes as
F3396/F3396M Practice for Aircraft Simplified Loads Cri-
defined in the F44 terminology standard. Use of the term
teria
airplane is used throughout this specification and will mean
2.2 U.S. Code of Federal Regulations:
“small airplane.”
14 CFR Part 23 Airworthiness Standards: Normal, Utility,
1.3 The applicant for a design approval must seek individual
Aerobatic and Commuter Category Airplanes (Amend-
guidance from their respective CAA body concerning the use
ment 62)
of this standard as part of a certification plan. For information
2.3 European Aviation Safety Agency Regulations:
on which CAA regulatory bodies have accepted this standard
Certification Specifications for Normal, Utility, Aerobatic,
(in whole or in part) as a means of compliance to their Small
and Commuter Category Aeroplanes (CS-23, Amendment
Airplane Airworthiness Rules (hereinafter referred to as “the
3)
Rules”), refer to ASTM F44 webpage (www.ASTM.org/
Certification Specifications for Very Light Aeroplanes (CS-
COMMITTEE/F44.htm) which includes CAA website links.
VLA, Amendment 1)
1.4 Units—Currently there is a mix of SI and Imperial units.
3. Terminology
In many locations, SI units have been included otherwise units
are as they appear in Amendment 62 of 14 CFR Part 23. In a
3.1 A listing of terms, abbreviations, acronyms, and sym-
future revision values will be consistently stated in SI units
bols related to aircraft covered by ASTM Committees F37 and
followed by Imperial units in square brackets. The values
F44 airworthiness design standards can be found in Terminol-
stated in each system may not be exact equivalents; therefore,
ogy F3060. Items listed below are more specific to this
each system shall be used independently of the other. Combin-
standard.
ing values from the two systems may result in non-
3.2 Definitions of Terms Specific to This Standard:
conformance with the standard.
3.2.1 chordwise, n—directed, moving, or placed along the
1.5 This standard does not purport to address all of the
chord of an airfoil section.
safety concerns, if any, associated with its use. It is the
3.2.2 downwash, n—the downward deflection of an air-
responsibility of the user of this standard to establish appro-
stream by an aircraft wing.
priate safety, health, and environmental practices and deter-
3.2.3 flight envelope, n—any combination of airspeed and
mine the applicability of regulatory limitations prior to use.
load factor on and within the boundaries of a flight envelope
1.6 This international standard was developed in accor-
that represents the envelope of the flight loading conditions
dance with internationally recognized principles on standard-
specified by the maneuvering and gust criteria.
ization established in the Decision on Principles for the
Development of International Standards, Guides and Recom- 3.2.4 flight load factor, n—represents the ratio of the aero-
mendations issued by the World Trade Organization Technical dynamic force component (acting normal to the assumed
Barriers to Trade (TBT) Committee.
For referenced ASTM standards, visit the ASTM website, www.astm.org, or
contact ASTM Customer Service at service@astm.org. For Annual Book of ASTM
This specification is under the jurisdiction of ASTM Committee F44 on General Standards volume information, refer to the standard’s Document Summary page on
Aviation Aircraft and is the direct responsibility of Subcommittee F44.30 on the ASTM website.
Structures. Available from U.S. Government Publishing Office (GPO), 732 N. Capitol St.,
Current edition approved Oct. 1, 2023. Published November 2023. Originally NW, Washington, DC 20401, http://www.gpo.gov.
approved in 2015. Last previous edition approved in 2023 as F3116/F3116M – 23. Available from European Aviation Safety Agency (EASA), Postfach 10 12 53,
DOI: 10.1520/F3116_F3116M-23A. D-50452 Cologne, Germany, https://www.easa.europa.eu/.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F3116/F3116M − 23a
longitudinal axis of the airplane) to the weight of the airplane. 4.3 Symmetrical Flight Conditions:
A positive flight load factor is one in which the aerodynamic 4.3.1 The appropriate balancing horizontal tail load must be
force acts upward, with respect to the airplane. accounted for in a rational or conservative manner when
determining the wing loads and linear inertia loads correspond-
3.2.5 propeller slipstream, n—the airstream pushed back by
ing to any of the symmetrical flight conditions specified in 4.4
a revolving aircraft propeller.
through 4.6.
3.2.6 spanwise, n—directed, moving, or placed along the
4.3.2 The incremental horizontal tail loads due to maneu-
span of an airfoil.
vering and gusts must be reacted by the angular inertia of the
3.2.7 winglet, n—a nearly vertical airfoil at an airplane’s
airplane in a rational or conservative manner.
wingtip.
4.3.3 Mutual influence of the aerodynamic surfaces must be
taken into account when determining flight loads.
3.3 Acronyms:
3.3.1 MCP—maximum continuous power
4.4 Flight Envelope:
4.4.1 General—Compliance with the strength requirements
3.4 Symbols:
of this subpart must be shown at any combination of airspeed
C = maximum airplane normal force coefficient
NA
and load factor on and within the boundaries of a flight
M = design cruising speed (Mach number)
C
envelope (similar to the one in 4.4.4) that represents the
V = design dive speed at zero or negative load factor
E
envelope of the flight loading conditions specified by the
V = stalling speed with flaps fully extended
SF
maneuvering and gust criteria of 4.4.2 and 4.4.3 respectively.
4. Flight Loads
4.4.2 Maneuvering Envelope—Except where limited by
maximum (static) lift coefficients, the airplane is assumed to be
4.1 Loads:
subjected to symmetrical maneuvers resulting in the following
4.1.1 Unless otherwise provided, prescribed loads are limit
limit load factors:
loads.
4.4.2.1 The positive maneuvering load factor specified in
4.1.2 Unless otherwise provided, the air, ground, and water
4.5 at speeds up to V ;
loads must be placed in equilibrium with inertia forces,
D
4.4.2.2 The negative maneuvering load factor specified in
considering each item of mass in the airplane. These loads must
4.5 at V ; and
be distributed to conservatively approximate or closely repre-
C
4.4.2.3 Factors varying linearly with speed from the speci-
sent actual conditions. Methods used to determine load inten-
fied value at V to 0.0 at V . For airplanes with a positive limit
sities and distribution on canard and tandem wing configura-
C D
maneuvering load factor greater than 3.8, use a value of –1.0 at
tions must be validated by flight test measurement unless the
V .
methods used for determining those loading conditions are
D
4.4.3 Gust Envelope:
shown to be reliable or conservative on the configuration under
4.4.3.1 The airplane is assumed to be subjected to sym-
consideration.
metrical vertical gusts in level flight. The resulting limit load
4.1.3 If deflections under load would significantly change
factors must correspond to the conditions determined as
the distribution of external or internal loads, this redistribution
follows:
must be taken into account.
(1) Positive (up) and negative (down) gusts of 15.24 m/s
4.1.4 Practice F3396/F3396M provides, within the limita-
[50 fps] at V must be considered at altitudes between sea level
tions specified within this practice, a simplified means of
C
and 6096 m [20 000 ft]. The gust velocity may be reduced
compliance with several of the requirements set forth in 4.2 to
linearly from 15.24 m/s [50 fps] at 6096 m [20 000 ft] to
4.26 and 7.1 to 7.9 that can be applied as one (but not the only)
7.62 m ⁄s [25 fps] at 15 240 m [50 000 ft]; and
means to comply.
(2) Positive and negative gusts of 7.62 m/s [25 fps] at V
D
4.2 General:
must be considered at altitudes between sea level and 6096 m
4.2.1 Flight load factors, n, represent the ratio of the
[20 000 ft]. The gust velocity may be reduced linearly from
aerodynamic force component (acting normal to the assumed
7.62 m/s [25 fps] at 6096 m [20 000 ft] to 3.81 m/s [12.5 fps]
longitudinal axis of the airplane) to the weight of the airplane.
at 15 240 m [50 000 ft].
A positive flight load factor is one in which the aerodynamic
(3) In addition, for level 4 airplanes, positive (up) and
force acts upward, with respect to the airplane.
negative (down) rough air gusts of 20.12 m/s [66 fps] at V
B
4.2.2 Compliance with the flight load requirements of this
must be considered at altitudes between sea level and 6096 m
subpart must be shown:
[20 000 ft]. The gust velocity may be reduced linearly from
4.2.2.1 At each critical altitude within the range in which
20.12 m/s [66 fps] at 6096 m [20 000 ft] to 11.58 m/s [38 fps]
the airplane may be expected to operate;
at 15 240 m [50 000 ft].
4.2.2.2 At each weight from the design minimum weight to
4.4.3.2 The following assumptions must be made:
the design maximum weight; and
(1) The shape of the gust is:
4.2.2.3 For each required altitude and weight, for any
U 2πs
practicable distribution of disposable load within the operating
de
U 5 1 2 cos (1)
S D
2 25C
limitations specified in 14 CFR Part 23, Sections 23.1583
through 23.1589.
where:
4.2.3 When significant, the effects of compressibility must
s = distance penetrated into gust (m or [ft]);
be taken into account.
F3116/F3116M − 23a
4.6.3 In the absence of a more rational analysis, the gust
C = mean geometric chord of wing (m or [ft]); and
load factors must be computed as follows:
U = derived gust velocity referred to in 4.4.3.1 (m/s or
de
[fps]).
K U Va
g de
n 5 11 (2)
(2) Gust load factors vary linearly with speed between V
C
W
S D
and V .
D
S
4.4.4 Flight Envelope—See Fig. 1.
where:
4.5 Limit Maneuvering Load Factors:
0.88μ
= gust alleviation factor;
g
K 5
g
4.5.1 The positive limit maneuvering load factor n may not
5.31μ
g
be less than:
2~W ⁄ S!
= airplane mass ratio;
μ 5
24,000 g
ρCag
4.5.1.1 2.1 1 , where W = design maximum take-
W110,000
U = derived gust velocities referred to in 4.4.3
de
off weight (lb), except that n need not be more than 3.8;
(fps).
4.5.1.2 6.0 for airplanes approved for aerobatics.
ρ = density of air (slugs/ft );
4.5.2 The negative limit maneuvering load factor may not
W/S = wing loading (psf) due to the applicable
be less than: weight of the airplane in the particular load
4.5.2.1 0.4 times the positive load factor; case;
C = mean geometric chord (ft);
4.5.2.2 0.5 times the positive load factor for airplanes
g = acceleration due to gravity (ft/s );
approved for aerobatics.
V = airplane equivalent speed (knots); and
4.5.3 Maneuvering load factors lower than those specified
a = slope of the airplane normal force coefficient
in this section may be used if the airplane has design features
curve C per radian if the gust loads are
NA
that make it impossible to exceed these values in flight.
applied to the wings and horizontal tail sur-
4.6 Gust Load Factors:
faces simultaneously by a rational method.
4.6.1 Each airplane must be designed to withstand loads on
The wing lift curve slope C per radian may
L
each lifting surface resulting from gusts specified in 4.4.3.
be used when the gust load is applied to the
4.6.2 The gust load factors for a canard or tandem wing
wings only and the horizontal tail gust loads
configuration must be computed using a rational analysis, or
are treated as a separate condition.
may be computed in accordance with 4.6.3, provided that the
resulting net loads are shown to be conservative with respect to
the gust criteria of 4.4.3.
NOTE 1—Point G need not be investigated when the supplementary condition specified in 4.14 is investigated.
FIG. 1 Flight Envelope
F3116/F3116M − 23a
4.7 Design Fuel Loads: 4.8.4.3 For the investigation of slipstream effects, the load
4.7.1 The disposable load combinations must include each factor may be assumed to be 1.0.
fuel load in the range from zero fuel to the selected maximum
4.9 Unsymmetrical Flight Conditions:
fuel load.
4.9.1 The airplane is assumed to be subjected to the unsym-
4.7.2 If fuel is carried in the wings, the maximum allowable
metrical flight conditions of 4.10 and 4.11. Unbalanced aero-
weight of the airplane without any fuel in the wing tank(s) must
dynamic moments about the center of gravity must be reacted
be established as “maximum zero wing fuel weight,” if it is less
in a rational or conservative manner, considering the principal
than the maximum weight.
masses furnishing the reacting inertia forces.
4.7.3 For level 4 airplanes, a structural reserve fuel
4.9.2 Airplanes approved for aerobatics must be designed
condition, not exceeding fuel necessary for 45 min of operation
for additional asymmetric loads acting on the wing and the
at maximum continuous power, may be selected. If a structural
horizontal tail.
reserve fuel condition is selected, it must be used as the
4.10 Rolling Conditions—The wing and wing bracing must
minimum fuel weight condition for showing compliance with
be designed for the following loading conditions:
the flight load requirements prescribed in this part and:
4.10.1 Unsymmetrical wing loads. Unless the following
4.7.3.1 The structure must be designed to withstand a
values result in unrealistic loads, the rolling accelerations may
condition of zero fuel in the wing at limit loads corresponding
be obtained by modifying the symmetrical flight conditions in
to:
4.4.4 as follows:
(1) 90 % of the maneuvering load factors defined in 4.5,
4.10.1.1 In Condition A, assume that 100 % of the semispan
and
wing airload acts on one side of the airplane and 70 % of this
(2) Gust velocities equal to 85 % of the values prescribed
load acts on the other side. For airplanes of more than 454 kg
in 4.4.3.
[1000 lb] design weight, the latter percentage may be increased
4.7.3.2 The fatigue evaluation of the structure must account
linearly with weight up to 75 % at 5670 kg [12 500 lb].
for any increase in operating stresses resulting from the design
4.10.1.2 For airplanes approved for aerobatics, in conditions
condition of 4.7.3.1.
A and F, assume that 100 % of the semispan wing airload acts
4.7.3.3 The flutter, deformation, and vibration requirements
on one side of the plane of symmetry and 60 % of this load acts
must also be met with zero fuel in the wings.
on the other side.
4.8 High Lift Devices:
4.10.2 The loads resulting from the aileron deflections and
4.8.1 If wing flaps or similar high lift devices are installed
speeds specified in 4.25, in combination with an airplane load
for use in take-off, approach, or landing, the airplane, with the
factor of at least two thirds of the positive maneuvering load
flaps fully deflected at V , is assumed to be subjected to
F
factor used for design. Unless the following values result in
symmetrical maneuvers and gusts resulting in limit load factors
unrealistic loads, the effect of aileron displacement on wing
within the range determined by:
torsion may be accounted for by adding the following incre-
4.8.1.1 Maneuvering, to a positive limit load factor of 2.0;
ment to the basic airfoil moment coefficient over the aileron
and
portion of the span in the critical condition determined in 4.4.4:
4.8.1.2 Positive and negative gust of 7.62 m ⁄s [25 fps]
∆c 5 20.01 δ (3)
m
acting normal to the flight path in level flight.
4.8.1.3 However, if an automatic flap load limiting device is
where:
used, the airplane may be designed for the critical combina-
Δc = the moment coefficient increment; and
m
tions of airspeed and flap position allowed by that device.
δ = the down aileron deflection in degrees in the critical
4.8.2 V must be assumed to be not less than 1.4 V or
F S condition.
1.8 V , whichever is greater, where:
SF
4.11 Yawing Conditions—The airplane must be designed for
4.8.2.1 V is the 1g computed stalling speed with flaps
S
yawing loads on the vertical surfaces resulting from the loads
retracted at the design weight; and
specified in 4.20 through 4.22.
4.8.2.2 V is the 1g computed stalling speed with flaps
SF
fully extended at the design weight. 4.12 Pressurized Cabin Loads—For each pressurized
compartment, the following applies:
4.8.3 In determining external loads on the airplane as a
4.12.1 The airplane structure must be strong enough to
whole, thrust, slipstream, and pitching acceleration may be
assumed to be zero. withstand the flight loads combined with pressure differential
loads from zero up to the maximum relief valve setting.
4.8.4 The flaps, their operating mechanism, and their sup-
porting structures, must be designed for the conditions pre- 4.12.2 The external pressure distribution in flight, and any
scribed in 4.8.1. In addition, with the flaps fully extended at V , stress concentrations, must be accounted for.
F
the following conditions, taken separately, must be accounted
4.12.3 If landings may be made with the cabin pressurized,
for: landing loads must be combined with pressure differential
4.8.4.1 A head-on gust having a velocity of 7.62 m/s loads from zero up to the maximum allowed during landing.
[25 fps] (EAS), combined with propeller slipstream corre-
4.12.4 The airplane structure must be strong enough to
sponding to 75 % of maximum continuous power; and withstand the pressure differential loads corresponding to the
4.8.4.2 The effects of propeller slipstream corresponding to maximum relief valve setting multiplied by a factor of 1.33,
maximum takeoff power. omitting other loads.
F3116/F3116M − 23a
4.12.5 If a pressurized cabin has two or more compartments 4.15.1 The airplane must be designed for the symmetrical
separated by bulkheads or a floor, the primary structure must be maneuvers and gusts prescribed in 4.4, 4.5, and 4.6, and the
designed for the effects of sudden release of pressure in any yawing maneuvers and lateral gusts in 4.20 and 4.21, with the
device extended at speeds up to the placard device extended
compartment with external doors or windows. This condition
speed; and
must be investigated for the effects of failure of the largest
opening in the compartment. The effects of intercompartmental 4.15.2 If the device has automatic operating or load limiting
features, the airplane must be designed for the maneuver and
venting may be considered.
gust conditions prescribed in 4.15.1 at the speeds and corre-
4.13 Unsymmetrical Loads Due to Engine Failure:
sponding device positions that the mechanism allows.
4.13.1 Multi-engine airplanes must be designed for the
4.16 Balancing Loads:
unsymmetrical loads resulting from the failure of the critical
4.16.1 A horizontal surface balancing load is a load neces-
engine including the following conditions in combination with
sary to maintain equilibrium in any specified flight condition
a single malfunction of the propeller drag limiting system,
with no pitching acceleration.
considering the probable pilot corrective action on the flight
4.16.2 Horizontal balancing surfaces must be designed for
controls:
the balancing loads occurring at any point on the limit
4.13.1.1 At speeds between V and V , the loads resulting
MC D
maneuvering envelope and in the flap conditions specified in
from power failure because of fuel flow interruption are
4.8.
considered to be limit loads.
4.16.3 For airplanes meeting the limitations of Practice
4.13.1.2 At speeds between V and V , the loads resulting
MC C
F3396/F3396M, Control Surface Loading (Level 1
from the disconnection of the engine compressor from the
Aeroplanes), the distribution of horizontal tail balancing loads,
turbine or from loss of the turbine blades are considered to be
Practice F3396/F3396M, Tail Surface Balancing and Maneu-
ultimate loads.
vering Load Distribution, may be used.
4.13.1.3 The time history of the thrust decay and drag
4.17 Maneuvering Loads for Horizontal Surfaces—Each
buildup occurring as a result of the prescribed engine failures
horizontal surface and its supporting structure, and the main
must be substantiated by test or other data applicable to the
wing of a canard or tandem wing configuration, if that surface
particular engine-propeller combination.
has pitch control, must be designed for the maneuvering loads
4.13.1.4 The timing and magnitude of the probable pilot
imposed by conditions 4.17.1 and 4.17.2. For airplanes meet-
corrective action must be conservatively estimated, consider-
ing the limitations of Practice F3396/F3396M, Control Surface
ing the characteristics of the particular engine-propeller-
Loading (Level 1 Aircraft), either the condition of 4.17.3 or
airplane combination.
4.17.4 can be used instead of the loads determined in condi-
4.13.2 Pilot corrective action may be assumed to be initiated
tions 4.17.1 and 4.17.2.
at the time maximum yawing velocity is reached, but not
4.17.1 A sudden movement of the pitching control at the
earlier than 2 s after the engine failure. The magnitude of the
speed V ,
A
corrective action may be based on the limit pilot forces
4.17.1.1 to the maximum aft movement (upward
specified in 7.4 except that lower forces may be assumed where
deflection), and
it is shown by analysis or test that these forces can control the
4.17.1.2 the maximum forward movement (downward
yaw and roll resulting from the prescribed engine failure
deflection), as limited by the control stops, or pilot effort,
conditions.
whichever is critical.
4.14 Rear Lift Truss:
4.17.1.3 For airplanes meeting the limitations of Practice
F3396/F3396M, Control Surface Loading (Level 1 Aeroplane),
4.14.1 If a rear lift truss is used, it must be designed for
the average loading of Practice F3396/F3396M, Acceptable
conditions of reversed airflow at a design speed of:
Methods for Control Surface Loads Calculations, Control
W
Surface Loads and the distribution for horizontal tail surfaces,
V 5 8.7Π18.7 ~knots! (4)
S
Practice F3396/F3396M, Tail Surface, Horizontal Down Load
Distribution, may be used.
where:
4.17.2 A sudden aft movement of the pitching control at
W/S = wing loading (lb/ft ) at design maximum takeoff
speeds above V , followed by a forward movement of the
A
weight.
pitching control resulting in the following combinations of
4.14.2 Either aerodynamic data for the particular wing
normal and angular acceleration:
section used, or a value of C equalling –0.8 with a chordwise
L
Normal Angular
Condition acceleration acceleration
distribution that is triangular between a peak at the trailing
(n) (radian/s )
edge and zero at the leading edge, must be used.
Nose-up pitching (down load) 1.0
1 n n 2 1.5
4.15 Speed Control Devices—If speed control devices (such s d
m m
V
as spoilers and drag flaps) are incorporated for use in enroute
Nose-down pitching (up load) n
m
2 n n 2 1.5
s d
m m
conditions:
V
F3116/F3116M − 23a
where:
Initial Final Load Factor
Speed
Condition Condition Increment
n = positive limit maneuvering load factor used in the
m
V A A n1 – 1
A 1
design of the airplane; and
A A 1 – n1
A G n4 – 1
V = initial speed in knots.
G A 1 – n4
V D D n2 – 1
4.17.2.1 The conditions in this section involve loads corre- D 1
D D 1 – n2
sponding to the loads that may occur in a “checked maneuver”
D E n3 – 1
(a maneuver in which the pitching control is suddenly dis-
E D 1 – n3
placed in one direction and then suddenly moved in the
4.17.4 For the purpose of this calculation, the difference in
opposite direction). The deflections and timing of the “checked
air speed between V and the value corresponding to point G
A
maneuver” must avoid exceeding the limit maneuvering load
on the maneuvering envelope can be ignored. The following
factor. The total horizontal surface load for both nose-up and
assumptions must be made:
nose-down pitching conditions is the sum of the balancing
4.17.4.1 The airplane is initially in level flight, and its
loads at V and the specified value of the normal load factor n,
attitude and airspeed do not change;
plus the maneuvering load increment due to the specified value
4.17.4.2 The loads are balanced by inertia forces;
of the angular acceleration. For airplanes meeting the limita-
tions of Practice F3396/F3396M, Control Surface Loading
4.17.4.3 The aerodynamic tail load increment is given by:
(Level 1 Aeroplanes), the maneuvering load increment in
X S a dϵ ρ S a l
cg ht ht 0 ht ht t
Practice F3396/F3396M, Maneuvering Tail Load Increment ∆P 5 ∆nMg 2 1 2 2 (5)
F S D S DG
l S a dα 2 M
t
(Up or Down) and; for Down Loads, the distributions for
where:
horizontal tail surfaces, Practice F3396/F3396M, Tail Surface,
Horizontal Down Load Distribution may be used. For Up ΔP = horizontal tail load increment, positive upwards (N),
Δn = load factor increment,
Loads, the distributions for vertical tail surfaces, Practice
M = mass of the airplane (kg),
F3396/F3396M, Tail Surface Vertical and Horizontal Up Load
g = acceleration due to gravity (m/s ),
Distribution may be used.
X = longitudinal distance of airplane c.g. aft of aerody-
cg
4.17.3 A sudden deflection of the elevator, the following
namic center of airplane less horizontal tail (m),
cases must be considered:
S = horizontal tail area (m ),
ht
4.17.3.1 Speed V , maximum upward deflection;
A
a = slope of horizontal tail lift curve per radian,
ht
dϵ
4.17.3.2 Speed V , maximum downward deflection; = rate of change of downwash angle with angle of
A
dα
attack,
4.17.3.3 Speed V , one-third maximum upward deflection;
D
ρ = density of air at sea-level (kg/m ),
4.17.3.4 Speed V , one-third maximum downward deflec-
D
l = tail arm (m),
t
tion.
S = wing area (m ), and
4.17.3.5 The following assumptions must be made:
a = slope of wing lift curve per radian.
(1) The airplane is initially in level flight, and its attitude
4.18 Gust Loads for Horizontal Surfaces:
and air speed do not change.
(2) The loads are balanced by inertia forces. 4.18.1 Each horizontal surface, other than a main wing,
must be designed for loads resulting from:
A sudden deflection of the elevator such as to cause the
normal acceleration to change from an initial value to a final 4.18.1.1 Gust velocities specified in 4.4.3 with flaps re-
value, the following cases being considered (see Fig. 2): tracted; and
FIG. 2 Pitching Maneuvers
F3116/F3116M − 23a
4.18.1.2 Positive and negative gusts of 7.62 m/s [25 fps] 4.20.1 At speeds up to V , the vertical surfaces must be
A
nominal intensity at V , corresponding to the flight conditions designed to withstand the following conditions. In computing
F
specified in 4.8.1.2.
the loads, the yawing velocity may be assumed to be zero:
4.18.2 For airplanes meeting the limitations of Practice 4.20.1.1 With the airplane in unaccelerated flight at zero
F3396/F3396M, Control Surface Loading (Level 1 yaw, it is assumed that the rudder control is suddenly displaced
Aeroplanes), the average loadings in Practice F3396/F3396M,
to the maximum deflection, as limited by the control stops or
Up and Down Gust Loading on Horizontal Tail Surface, and by limit pilot forces.
the distributions for vertical tail surfaces, Practice F3396/
4.20.1.2 With the rudder deflected as specified in 4.20.1.1, it
F3396M, Tail Surface Vertical and Horizontal Up Load
is assumed that the airplane yaws to the overswing sideslip
Distribution, may be used to determine the incremental gust
angle. In lieu of a rational analysis, an overswing angle may be
loads for the requirements of 4.18.1 applied as both up and
assumed equal to 1.5 times the static sideslip angle of 4.20.1.3.
down increments for 4.18.3.
4.20.1.3 A yaw angle of 15° with the rudder control
4.18.3 When determining the total load on the horizontal
maintained in the neutral position (except as limited by pilot
surfaces for the conditions specified in 4.18.1, the initial
strength).
balancing loads for steady unaccelerated flight at the pertinent
4.20.2 For airplanes meeting the limitations of Practice
design speeds V , V , and V must first be determined. The
F C D
F3396/F3396M, Control Surface Loading (Level 1
incremental load resulting from the gusts must be added to the
Aeroplanes), the average loading of Practice F3396/F3396M,
initial balancing load to obtain the total load.
Limit Average Maneuvering Control Surface Loading and the
4.18.4 In the absence of a more rational analysis, the
distributions in Practice F3396/F3396M, Tail Balancing and
incremental load due to the gust must be computed as follows
Maneuver Load Distribution, Tail Surface, Horizontal Down
only on airplane configurations with aft-mounted, horizontal
Load Distribution, Tail Surface, Vertical and Horizontal Up
surfaces, unless its use elsewhere is shown to be conservative:
Load Distribution, may be used instead of the requirements of
4.20.1.2, 4.20.1.1, and 4.20.1.3, respectively.
K U Va S d
g de ht ht e
∆L 5 1 2 (6)
S D
ht
498 d
} 4.20.3 For level 4 airplanes, the loads imposed by the
following additional maneuver must be substantiated at speeds
where:
from V to V /M . When computing the tail loads:
A D D
ΔL = incremental horizontal tail load (lb);
ht
4.20.3.1 The airplane must be yawed to the largest attain-
K = gust alleviation factor defined in 4.6;
g
able steady state sideslip angle, with the rudder at maximum
U = derived gust velocity (fps);
de
deflection caused by any one of the following:
V = airplane equivalent speed (knots);
(1) Control surface stops;
a = slope of aft horizontal tail lift curve (per radian);
ht
S = area of aft horizontal lift surface (ft ); and
(2) Maximum available booster effort;
ht
d
= downwash factor.
e (3) Maximum pilot rudder force as shown in Fig. 3.
1 2
S D
d
}
4.20.3.2 The rudder must be suddenly displaced from the
4.19 Unsymmetrical Loads:
maximum deflection to the neutral position.
4.19.1 Horizontal surfaces other than main wing and their
4.20.4 The yaw angles specified in 4.20.1.3 may be reduced
supporting structure must be designed for unsymmetrical loads
if the yaw angle chosen for a particular speed cannot be
arising from yawing and slip-stream effects, in combination
exceeded in:
with the loads prescribed for the flight conditions set forth in
4.20.4.1 Steady slip conditions;
4.16 through 4.18.
4.20.4.2 Uncoordinated rolls from steep banks; or
4.19.2 In the absence of more rational data for airplanes that
4.20.4.3 For multi-engine airplanes, the sudden failure of
are conventional in regard to location of engines, wings,
the critical engine with delayed corrective action.
horizontal surfaces other than main wing, and fuselage shape:
4.19.2.1 100 % of the maximum loading from the symmetri- 4.21 Gust Loads for Vertical Surfaces:
cal flight conditions may be assumed on the surface on one side
4.21.1 Vertical surfaces must be designed to withstand, in
of the plane of symmetry; and
unaccelerated flight at speed V , lateral gusts or the values
C
4.19.2.2 The following percentage of that loading must be prescribed for V in 4.4.3.
C
applied to the opposite side: Percent = 100 – 10 (n – 1), where
4.21.2 In addition, for level 4 airplanes, the airplane is
n is the specified positive maneuvering load factor, but this
assumed to encounter derived gusts normal to the plane of
value may not be more than 80 %.
symmetry while in unaccelerated flight at V , V , V , and V .
B C D F
4.19.3 For airplanes that are not conventional (such as The derived gusts and airplane speeds corresponding to these
airplanes with horizontal surfaces other than main wing having conditions, as determined by 4.6 and 4.8, must be investigated.
appreciable dihedral or supported by the vertical tail surfaces) The shape of the gust must be as specified in 4.4.3.2(1).
the surfaces and supporting structures must be designed for
4.21.3 In the absence of a more rational analysis, the gust
combined vertical and horizontal surface loads resulting from
load must be computed as follows:
each prescribed flight condition taken separately.
K U Va S
gt de vt vt
L 5 (7)
vt
4.20 Maneuvering Loads for Vertical Surfaces: 498
F3116/F3116M − 23a
FIG. 3 Maximum Pilot Rudder Force
where: 4.22.2 If outboard fins or winglets extend above and below
the horizontal surface, the critical vertical surface loading (the
L = vertical surface loads (lb);
vt
0.88μ
load per unit area as determined under 4.20 and 4.21) must be
g = gust alleviation factor;
K 5
g
5.31μ applied to:
gt
2W K
= lateral mass ratio; 4.22.2.1 The part of the vertical surfaces above the horizon-
μ 5
S D
gt
ρc¯ ga S l
t vt vt vt tal surface with 80 % of that loading applied to the part below
U = derived gust velocity (fps);
the horizontal surface; and
de
ρ = air density (slugs/ft );
4.22.2.2 The part of the vertical surfaces below the horizon-
W = the applicable weight of the airplane
tal surface with 80 % of that loading applied to the part above
in the particular load case (lb);
the horizontal surface.
S = area of vertical surface (ft );
vt
4.22.3 The end plate effects of outboard fins or winglets
c¯ = mean geometric chord of vertical sur-
t
must be taken into account in applying the yawing conditions
face (ft);
of 4.20 and 4.21 to vertical surfaces in 4.22.2.
a = lift curve slope of vertical surface
vt
4.22.4 When rational methods are used for computing loads,
(per radian);
the maneuvering loads of 4.20 on the vertical surfaces and the
K = radius of gyration in yaw (ft);
one-g horizontal surface load, including induced loads on the
l = distance from airplane c.g. to lift
vt
horizontal surface and moments or forces exerted on the
center of vertical surface (ft);
horizontal surfaces by the vertical surfaces, must be applied
g = acceleration due to gravity (ft/s );
simultaneously for the structural loading condition.
and
V = equivalent airspeed (knots).
4.23 Combined Loads on Tail Surfaces (for airplanes meet-
ing the limitations of Practice F3396/F3396M, Control Surface
4.21.4 For airplanes meeting the limitations of Practice
Loading (Level 1 Aeroplanes)):
F3396/F3396M, Control Surface Loading (Level 1
4.23.1 With the airplane in a loading condition correspond-
Aeroplanes), the average loading in Practice F3396/F3396M,
ing to point A or D in the V-n diagram (whichever condition
Gust Loading on Vertical Tail Surface, and the distribution in
leads to the higher balance load) the loads on the horizontal tail
Practice F3396/F3396M, Tail Surface, Vertical and Horizontal
must be combined with those on the vertical tail as specified in
Up Load Distribution, may be used.
4.20.
4.22 Outboard Fins or Winglets:
4.23.2 75 % of the loads according to 4.17 for the horizontal
4.22.1 If outboard fins or winglets are included on the
tail and 4.20 for the vertical tail must be assumed to be acting
horizontal surfaces or wings, the horizontal surfaces or wings
simultaneously.
must be designed for their maximum load in combination with
loads induced by the fins or winglets and moments or forces 4.24 Additional Loads Applicable to V-tails—(for airplanes
exerted on the horizontal surfaces or wings by the fins or meeting the limitations of Practice F3396/F3396M, Control
winglets. Surface Loading (Level 1 Aeroplanes))—An airplane with
F3116/F3116M − 23a
V-tail must be designed for a gust acting perpendicularly with 5.1.2.4 Compliance with 5.1.2.1 and 5.1.2.2 need not be
respect to one of the tail surfaces at speed V . This case is shown if V /M is selected so that the minimum speed margin
E D D
supplemental to the equivalent horizontal and vertical tail cases between V /M and V /M is the greater of the following:
C C D D
specified. Mutual interference between the V-tail surfaces must (1) The speed increase resulting when, from the initial
be adequately accounted for. condition of stabilized flight at V /M , the airplane is assumed
C C
to be upset, flown for 20 s along a flight path 7.5° below the
4.25 Ailerons:
initial path, and then pulled up with a load factor of 1.5 (0.5 g
4.25.1 The ailerons must be designed for the loads to which
acceleration increment). At least 75 % maximum continuous
they are subjected:
power for reciprocating engines, and maximum cruising power
4.25.1.1 In the neutral position during symmetrical flight
for turbines, or, if less, the power required for V /M for both
C C
conditions; and
kinds of engines, must be assumed until the pullup is initiated,
4.25.1.2 By the following deflections (except as limited by
at which point power reduction and pilot-controlled drag
pilot effort), during unsymmetrical flight conditions:
devices may be used; and either:
(1) Sudden maximum displacement of the aileron control
(2) Mach 0.05 (at altitudes where M is established); or
D
at V . Suitable allowance may be made for control system
(3) Mach 0.07 for level 4 airplanes (at altitudes where M
A
D
deflections.
is established) unless a rational analysis, including the effects
(2) Sufficient deflection at V , where V is more than V , to
of automatic systems, is used to determine a lower margin. If
C C A
produce a rate of roll not less than obtained in 4.25.1.2.
a rational analysis is used, the minimum speed margin must be
(3) Sufficient deflection at V to produce a rate of roll not
enough to provide for atmospheric variations (such as horizon-
D
less than one-third of that obtained in 4.25.1.2.
tal gusts), and the penetration of jet streams or cold fronts),
4.25.2 For airplanes meeting the limitations of Practice instrument errors, airframe production variations, and must not
F3396/F3396M, Control Surface Loading (Level 1 be less than Mach 0.05.
Aeroplanes), the average loading of Practice F3396/F3396M, 5.1.3 Design Maneuvering Speed V —For V , the following
A A
Control Surface Loading (Level 1 Aeroplanes), Control Sur- applies:
face Loads, and in Practice F3396/F3396M, The Limit Average
5.1.3.1 V may not be less than V =n where:
A S
Maneuvering Control Surface Loading, and the distribution in
(1) V is a 1g computed stalling speed with flaps retracted
S
Practice F3396/F3396M, Aileron Load Distribution, may be
(normally based on the maximum airplane normal force
used.
coefficients, C ) at either (1) the particular weight under
NA
consideration or (2) the design maximum takeoff weight; and
4.26 Special Devices—The loading for special devices using
(2) n is the limit maneuvering load factor used in design.
aerodynamic surfaces (such as slots, slats, and spoilers) must
be determined from test data. 5.1.3.2 The value of V need not exceed the value of V
A C
used in design.
5. Design Airspeeds 5.1.4 Design Speed for Maximum Gust Intensity, V —For
B
V , the following apply:
B
5.1 Design Airspeeds—Except as provided in 5.1.1.4, the
5.1.4.1 V may not be less than the speed determined by the
B
selected design airspeeds are equivalent airspeeds (EAS).
intersection of the line representing the maximum positive lift,
5.1.1 Design Cruising Speed, V —For V , the following
C C
C , and the line representing the rough air gust velocity on
NMAX
apply:
the gust V-n diagram, or V =n , whichever is less, where:
S g
5.1.1.1 Where W/S = wing loading at the design maximum
(1) n is the positive airplane gust load factor due to gust,
g
takeoff weight (lb/ft ), V (in knots) may not be less than:
C
at speed V (in accordance with 4.6), and at the particular
C
(1) 33=W⁄S; and
weight under consideration; and
(2) 36=W⁄S (for airplanes approved for aerobatics).
(2) V is the 1g stalling speed with the flaps retracted at the
S
particular weight under consideration.
5.1.1.2 For values of W/S more than 20, the multiplying
factors may be decreased linearly with W/S to a value of 28.6 5.1.4.2 V need not be greater than V .
B C
where W/S = 100.
6. Engine Mount Loads
5.1.1.3 V need not be more than 0.9 V at sea level.
C H
6.1 Engine Torque:
5.1.1.4 At altitudes where an M is established, a cruising
D
6.1.1 Each engine mount and its supporting structure must
speed M limited by compressibility may be selected.
C
be designed for the effects of:
5.1.2 Design Dive Speed, V —For V , the following apply:
D D
6.1.1.1 A limit engine torque corresponding to takeoff
5.1.2.1 V /M may not be less than 1.25 V /M : and
D D C C
power and, if applicable, propeller speed acting simultaneously
5.1.2.2 With V min, the required minimum design cruising
with 75 % of the limit loads from flight condition A of 4.4.4;
C
speed, V (in knots) may not be less than:
D 6.1.1.2 The limit engine torque as specified in 6.1.3 acting
(1) 1.40 V min; and
simultaneously with the limit loads from flight condition A of
C
(2) 1.55 V min (for airplanes approved for aerobatics).
C 4.4.4; and
5.1.2.3 For values of W/S more than 20, the multiplying 6.1.1.3 For turbo-propeller installations, in addition to the
factors in 5.1.2.2 may be decreased linearly with W/S to a value conditions specified in 6.1.1.1 and 6.1.1.2, a limit engine
of 1.35 where W/S = 100. torque corresponding to takeoff power and propeller speed,
F3116/F3116M − 23a
multiplied by a factor accounting for propeller control system 7.1.2 For airplanes meeting the limitations of Practice
malfunction, including quick feathering, acting simultaneously F3396/F3396M, Control Surface Loading (Level 1 Aero-
with 1g level flight loads. In the absence of a rational analysis, planes) and if allowed by the following paragraphs, the values
a factor of 1.6 must be used. of control surface loading in Practice F3396/F3396M, Control
Surface Loading (Level 1 Aeroplanes), may be used to deter-
6.1.2 For turbine engine installations, the engine mounts
and supporting structure must be designed to withstand each of mine the detailed rational requirements of 4.16 through 4.26
and 7.4 through 7.9, unless these values result in unrealistic
the following:
6.1.2.1 A limit engine torque load imposed by sudden loads.
engine stoppage due to malfunction or structural failure (such
7.2 Loads Parallel to Hinge Line:
as compressor jamming).
7.2.1 Control surfaces and supporting hinge brackets must
6.1.2.2 A limit engine torque load imposed by the maximum
be designed to withstand inertial loads acting parallel to the
acceleration of the engine.
hinge line.
6.1.3 The limit engine torque to be considered under 6.1.1
7.2.2 In the absence of more rational data, the inertia loads
must be obtained by multiplying the mean torque for maximum
may be assumed to be equal to KW, where:
continuous power by a factor determined as follows:
7.2.2.1 K = 24 for vertical surfaces;
6.1.3.1 1.25 for turbo-propeller installations;
7.2.2.2 K = 12 for horizontal surfaces; and
6.1.3.2 For four-stroke engines:
7.2.2.3 W = weight of the movable surfaces.
(1) 1.33 for engines with five or more cylinders,
7.3 Control System Loads:
(2) 2, 3, 4, or 8 for engines with four, three, two, or one
7.3.1 Each flight control system and its supporting structure
cylinders, respectively.
must be designed for loads corresponding to at least 125 % of
6.1.3.3 For two-stroke engines:
the computed hinge moments of the movable control surface in
(1) 2 for engines with three or more cylinders,
the conditions prescribed in 4.16 through 4.26 and 7.1 through
(2) 3 or 6, for engines with two or one cylinders respec-
7.9. In addition, the following apply:
tively.
7.3.1.1 The system limit loads need not exceed the higher of
6.2 Side Load on Engine Mount:
the loads that can be produced by the pilot and automatic
6.2.1 Each engine mount and its supporting structure must
devices operating the controls. However, autopilot forces need
be designed for a limit load factor in a lateral direction, for the
not be added to pilot forces. The system must be designed for
side load on the engine mount, of not less than:
the maximum effort of the pilot or autopilot, whichever is
6.2.1.1 1.33, or
higher. In add
...