ASTM F3396/F3396M-23a

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it may not be technically possible to adequately depict all changes accurately, ASTM recommends that users consult prior editions as appropriate. In all cases only the current version
of the standard as published by ASTM is to be considered the official document.
Designation: F3396/F3396M − 23 F3396/F3396M − 23a
Standard Practice for
Aircraft Simplified Loads Criteria
This standard is issued under the fixed designation F3396/F3396M; 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
1.1 This practice provides an acceptable means of meeting the airworthiness requirements for the flight design loads and
conditions of small normal category level 1 and 2 aeroplanes. The material was developed through open consensus of international
experts in general aviation. This information was created by focusing on Normal Category aeroplanes. The content may be more
broadly applicable; it is the responsibility of the applicant to substantiate broader applicability as a specific means of compliance.
The topics covered within this practice are: Simplified Design Load Criteria, Acceptable Methods for Control Surface Loads
Calculations, Acceptable Methods for Primary Control System Loads Calculations, and Control Surface Loading (Level 1
Aeroplanes).
1.2 This practice is applicable to normal category, low-speed, level 1 and 2 aeroplanes. Use of the term aeroplane used throughout
this practice will mean “normal category, low-speed, level 1 or 2 aeroplane,” unless otherwise stated.
1.3 An applicant intending to propose this information as means of compliance for a design approval must seek guidance from
their respective oversight authority (for example, published guidance from applicable CAAs) concerning the acceptable use and
application thereof. For information on which oversight authorities have accepted this standard (in whole or in part) as an
acceptable means of compliance to their regulatory requirements (hereinafter “the Rules”), refer to the ASTM Committee F44 web
page (www.astm.org/COMMITTEE/F44.htm).
1.4 This document may present information in either SI units, English Engineering units, or both. The values stated in each system
are not necessarily exact equivalents; therefore, to ensure conformance with the standard, each system shall be used independently
of the other, and values from the two systems shall not be combined.
1.5 This standard does not purport to address all of the safety concerns, if any, associated with its use. It is the responsibility
of the user of this standard to establish appropriate safety, health, and environmental practices and determine the applicability of
regulatory limitations prior to use.
1.6 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.
2. Referenced Documents
2.1 ASTM Standards:
F3060 Terminology for Aircraft
This practice is under the jurisdiction of ASTM Committee F44 on General Aviation Aircraft and is the direct responsibility of Subcommittee F44.30 on Structures.
Current edition approved March 15, 2023Oct. 1, 2023. Published June 2023November 2023. Originally approved in 2020. Last previous edition approved in 20202023
as F3396/F3396M – 20.F3396/F3396M – 23. DOI: 10.1520/F3396_F3396M-23.10.1520/F3396_F3396M-23A.
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 Standards
volume information, refer to the standard’s Document Summary page on the ASTM website.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F3396/F3396M − 23a
F3116/F3116M Specification for Design Loads and Conditions
2.2 U.S. Code of Federal Regulations:
14 CFR Part 23 Airworthiness Standards: Normal, Utility, Aerobatic and Commuter Category Airplanes (Amendment 62)
2.3 European Aviation Safety Agency (EASA) Regulations:
CS-23, Amendment 4 Certification Specifications for Normal, Utility, Aerobatic, and Commuter Category Aeroplanes
CS-VLA, Amendment 1 Certification Specifications for Very Light Aeroplanes
3. Terminology
3.1 A listing of terms, abbreviations, acronyms, and symbols related to aircraft covered by ASTM Committees F37 and F44
airworthiness design standards can be found in Terminology F3060. Items listed below are more specific to this standard.
3.2 Definitions of Terms Specific to This Standard:
3.2.1 chordwise, n—directed, moving, or placed along the chord of an airfoil section.
3.2.2 downwash, n—the downward deflection of an air-stream by an aircraft wing.
3.2.3 flight envelope, n—any combination of airspeed and load factor on and within the boundaries of a flight envelope that
represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria.
3.2.4 flight load factor, n—represents the ratio of the aero-dynamic force component (acting normal to the assumed longitudinal
axis of the aeroplane) to the weight of the aeroplane. A positive flight load factor is one in which the aerodynamic force acts
upward, with respect to the aeroplane.
3.2.5 propeller slipstream, n—the airstream pushed back by a revolving aircraft propeller.
3.2.6 spanwise, n—directed, moving, or placed along the span of an airfoil.
3.2.7 winglet, n—a nearly vertical airfoil at an aeroplane’s wingtip.
3.3 Symbols and Abbreviations:
3.3.1 C —maximum aeroplane normal force coefficient
NA
3.3.2 M —design cruising speed (Mach number)
C
3.3.3 MCP—maximum continuous power
3.3.4 n —aeroplane positive maneuvering limit load factor
3.3.5 n —aeroplane negative maneuvering limit load factor
3.3.6 n —aeroplane positive gust limit load factor at V
3 C
3.3.7 n —aeroplane negative gust limit load factor at V
4 C
3.3.8 n —aeroplane positive limit load factor with flaps fully extended at V
flap F
3.3.9 V —minimum design maneuvering speed = 15.0=n W⁄S knots (however this need not exceed V used in design)
A min 1 C
=
3.3.10 V —minimum design cruising speed = 17.0 n W⁄S knots (however this need not exceed 0.9V , see 5.2.5.2)
C min 1 H
Available from U.S. Government Publishing Office (GPO), 732 N. Capitol St., NW, Washington, DC 20401, http://www.gpo.gov.
Available from the European Aviation Safety Agency (EASA), Postfach 10 12 53, D-50452 Koeln, Germany, https://www.easa.europa.eu/.
F3396/F3396M − 23a
3.3.11 V —minimum design dive speed = 24.0=n W⁄S knots (however this need not exceed 1.4V =n ⁄3.8)
D min 1 Cmin 1
3.3.12 V —design dive speed at zero or negative load factor
E
3.3.13 V —minimum design flap speed = 11.0=n W⁄S knots
F min 1
3.3.14 V —stalling speed with flaps fully extended
SF
4. Significance and Use
4.1 This practice provides one means for determining the aeroplane structural loads for flight, control surfaces, and control
systems. This practice satisfies the simplified loads requirements set forth in Specification F3116/F3116M for Normal Category
Aeroplanes.
5. Simplified Design Load Criteria
5.1 Limitations:
5.1.1 The methods provided in this section provide one possible means (but not the only possible means) of compliance and can
only be applied to Normal Category, low-speed, level 1 and level 2 aeroplanes.
5.1.2 These methods may be applied to aeroplanes meeting the following limitations without further justification:
5.1.2.1 A single engine excluding turbine powerplants.
5.1.2.2 A main wing located closer to the aeroplane’s center of gravity than to the aft, fuselage-mounted, empennage.
5.1.2.3 A main wing that contains a quarter-chord sweep angle of not more than 15° fore or aft.
5.1.2.4 A main wing that is equipped with trailing-edge controls (ailerons or flaps, or both).
5.1.2.5 A main wing aspect ratio not greater than 7.0.
5.1.2.6 A main wing that does not have winglets, outboard fins, or other wingtip devices.
5.1.2.7 A horizontal tail aspect ratio not greater than 4.0.
5.1.2.8 A horizontal tail volume coefficient not less than 0.34.
5.1.2.9 A vertical tail aspect ratio not greater than 2.0.
5.1.2.10 A vertical tail planform area not greater than 10 % of the wing planform area.
5.1.2.11 Horizontal and vertical tail airfoil sections must both be symmetrical.
5.1.3 This section may be used outside of the limitations in 5.1.2 when evidence can be provided that the method provides safe
and reliable results.
5.1.4 Aeroplanes with any of the following design features shall not use this section:
5.1.4.1 Canard, tandem-wing, close-coupled, or tailless arrangements of the lifting surfaces;
5.1.4.2 Biplane or multiplane wing arrangements;
5.1.4.3 V-tail or any arrangement where the horizontal stabilizer is supported by the vertical stabilizer (T-tail, cruciform-tail (+),
etc.);
F3396/F3396M − 23a
5.1.4.4 Wings with slatted lifting surfaces; and
5.1.4.5 Full-flying stabilizing surfaces (horizontal and vertical).
5.2 Flight Loads:
5.2.1 Each flight load may be considered independent of altitude and, except for the local supporting structure for dead weight
items, only the maximum design weight conditions must be investigated.
5.2.2 Table 1 must be used to determine values of n ,n ,n , and n , corresponding to the maximum design weights. Fig. 1 presents
1 2 3 4
a generalized flight envelope.
5.2.3 Fig. 2 and Fig. 3 must be used to determine values of n and n , corresponding to the minimum flying weights, and, if these
3 4
load factors are greater than the load factors at the design weight, the supporting structure for dead weight items must be
substantiated for the resulting higher load factors.
5.2.4 Each specified wing and tail loading is independent of the center of gravity range. However, a center of gravity (c.g.) range
must be selected for the aeroplane and the basic fuselage structure must be investigated for the most adverse dead weight loading
conditions for the c.g. range selected.
5.2.5 The following loads and loading conditions are the minimums for which strength must be provided in the structure:
5.2.5.1 Aeroplane Equilibrium—The aerodynamic wing loads may be considered to act normal to the relative wind, and to have
a magnitude of 1.05 times the aeroplane normal loads (as determined from 5.3.2 and 5.3.3) for the positive flight conditions and
a magnitude equal to the aeroplane normal loads for the negative conditions. Each chordwise and normal component of this wing
load must be considered.
5.2.5.2 Minimum Design Airspeeds—The minimum design airspeeds may not be less than the minimum speeds found in 3.3. In
addition, V need not exceed values of 0.9V actually obtained at sea level for the lowest design weight for which certification
C min H
is desired. In computing these minimum design airspeeds, n may not be less than 3.8.
5.2.5.3 Flight Load Factor—The limit flight load factors specified in Table 1 represent the ratio of the aerodynamic force
component (acting normal to the assumed longitudinal axis of the aeroplane) to the weight of the aeroplane. A positive flight load
factor is an aerodynamic force acting upward, with respect to the aeroplane.
5.3 Flight Conditions:
5.3.1 General—Each design condition in 5.3.2 and 5.3.3 must be used to assure sufficient strength for each condition of speed and
load factor on or within the boundary of a V-n diagram for the aeroplane similar to the diagram in Fig. 1. This diagram must also
be used to determine the aeroplane structural operating limitations as specified in 14 CFR Part 23, Sec. 23.1501 (c) through
23.1513 and 23.1519.
5.3.2 Symmetrical Flight Conditions—The aeroplane must be designed for symmetrical flight conditions as follows:
5.3.2.1 The aeroplane must be designed for at least the four basic flight conditions, “A,” “D,” “E,” and “G” as noted on the flight
envelope of Fig. 1. In addition, the following requirements apply:
TABLE 1 Minimum Design Limit Flight Load Factors
Flight Load Factors Not Approved for Approved for
Aerobatics Aerobatics
n 3.8 6.0
n –0.5 n
2 1
Flaps Up
n Find from Fig. 2
n Find from Fig. 3
n 0.5 n
flap 1
Flaps Down
A
n Zero
flap
A
Vertical wing load may be assumed equal to zero and only the flap part of the
wing need be checked for this condition.
F3396/F3396M − 23a
NOTE 1—Conditions “C” and “F” of Fig. 1 need only be investigated when n W/Sor n W/Sare greater than n W/Sand n W/S, respectively.
3 4 1 2
NOTE 2—Condition “G” need not be investigated when the supplementary condition specified for a rear lift truss is investigated.
FIG. 1 Generalized Flight Envelope
FIG. 2 Chart for Finding n Factor at Speed V
3 C
(1) The design limit flight load factors corresponding to conditions “D” and “E” of Fig. 1 must be at least as great as those
specified in Table 1 and Fig. 1, and the design speed for these conditions must be at least equal to the value of V from 3.3.
D min
(2) For conditions “A” and “G” of Fig. 1, the load factors must correspond to those specified in Table 1, and the design speeds
must be computed using these load factors with the maximum static lift coefficient C determined by the applicant. However, in
NA
the absence of more precise computations, these latter conditions may be based on a value of C = 6 1.35 and the design speed
NA
for condition “A” may be less than V .
A min
(3) Conditions “C” and “F” of Fig. 1 need only be investigated when n W/Sor n W/Sare greater than n W/Sor n W/S,
3 4 1 2
respectively.
5.3.2.2 If flaps or other high lift devices intended for use at the relatively low airspeed of approach, landing, and takeoff, are
installed, the aeroplane must be designed for the two flight conditions corresponding to the values of limit flap-down factors
specified in Table 1 with the flaps fully extended at not less than the design flap speed V from 3.3.
F min
5.3.3 Unsymmetrical Flight Conditions—Each affected structure must be designed for unsymmetrical loadings as follows:
5.3.3.1 The aft fuselage-to-wing attachment must be designed for the critical vertical surface load determined in accordance with
6.2.3.1 and 6.2.3.2.
5.3.3.2 The wing and wing carry-through structures must be designed for 100 % of condition “A” loading on one side of the plane
of symmetry and 70 % on the opposite side, or 60 % on the opposite side for aeroplanes approved for aerobatics.
F3396/F3396M − 23a
FIG. 3 Chart for Finding n Factor at Speed V
4 C
5.3.3.3 The wing and wing carry-through structures must be designed for the loads resulting from a combination of 75 % of the
positive maneuvering wing loading on both sides of the plane of symmetry and the maximum wing torsion resulting from aileron
displacement. The effect of aileron displacement on wing torsion at V or V using the basic airfoil moment coefficient, C ,
C A mo
modified over the aileron portion of the span, must be computed as follows:
(1) C = C 6 0.01 δ (up aileron side) wing basic airfoil.
m mo u
(2) C = C 6 0.01 δ (down aileron side) wing basic airfoil, where δ is the up aileron deflection and δ is the down aileron
m mo d u d
deflection.
5.3.3.4 ∆ , which is the sum of δ + δ , must be computed as follows:
critical u d
(1) Computeandfrom the formulas:
a b
V
A
∆ 5 ×∆ (1)
a P
V
C
V
A
∆ 5 0.5 ×∆ (2)
b P
V
D
where:
∆ = the maximum total deflection (sum of both aileron deflections) at V with V ,V , and V described in 5.2.5.2.
p A A C D
(2) Compute K from the formula:
C 2 0.01δ V
~ !
mo b D
K 5 (3)
~C 2 0.01δ !V
mo a C
where:
δ = the down aileron deflection corresponding to ∆ , and
a a
δ = the down aileron deflection corresponding to ∆ as computed in 5.3.3.4(1).
b b
(3) If K is less than 1.0,isand must be used to determine δ and δ . In this case, V is the critical speed which must
a critical u d C
be used in computing the wing torsion loads over the aileron span.
(4) If K is equal to or greater than 1.0,isand must be used to determine δ and δ . In this case, V is the critical
b critical u d D
speed which must be used in computing the wing torsion loads over the aileron span.
5.3.4 Supplementary Conditions: Rear Lift Truss; Engine Torque; Side Load on Engine Mount—Each of the following
supplementary conditions must be investigated:
5.3.4.1 In designing the rear lift truss, the special condition specified in Specification F3116/F3116M may be investigated instead
of condition “G” of Fig. 1.
F3396/F3396M − 23a
5.3.4.2 Each engine mount and its supporting structures must be designed for:
(1) The maximum limit torque corresponding to maximum take-off power (MTO power) and propeller speed acting
simultaneously with 75 % of the limit loads resulting from the maximum positive maneuvering flight load factor n .
(2) The maximum limit t
...