ASTM F2245-23

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: F2245 − 23
Standard Specification for
Design and Performance of a Light Sport Airplane
This standard is issued under the fixed designation F2245; 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 F2316 Specification for Airframe Emergency Parachutes
F2339 Practice for Design and Manufacture of Reciprocat-
1.1 This specification covers airworthiness requirements for
ing Spark Ignition Engines for Light Sport Aircraft
the design of powered fixed wing light sport aircraft, an
F2483 Practice for Maintenance and the Development of
“airplane.”
Maintenance Manuals for Light Sport Aircraft
1.2 This specification is applicable to the design of a light
F2506 Specification for Design and Testing of Light Sport
sport aircraft/airplane as defined by regulations and limited to
Aircraft Propellers
VFR flight.
F2538 Practice for Design and Manufacture of Reciprocat-
1.3 Units—The values given in this standard are in SI units ing Compression Ignition Engines for Light Sport Aircraft
F2564 Specification for Design and Performance of a Light
and are to be regarded as standard. The values given in
parentheses are mathematical conversions to inch-pound (or Sport Glider
F2746 Specification for Pilot’s Operating Handbook (POH)
other) units that are provided for information only and are not
considered standard. The values stated in each system may not for Light Sport Airplane
F2840 Practice for Design and Manufacture of Electric
be exact equivalents. Where it may not be clear, some
equations provide the units of the result directly following the Propulsion Units for Light Sport Aircraft
F3619 Specification for Aeroelasticity Requirements for a
equation.
Light Sport Airplane
1.4 This standard does not purport to address all of the
2.2 Federal Aviation Regulations and Advisory Circulars:
safety concerns, if any, associated with its use. It is the
responsibility of the user of this standard to establish appro- 14 CFR Part 33 Airworthiness Standards: Aircraft Engines
14 CFR Part 35 Airworthiness Standards: Propellers
priate safety, health, and environmental practices and deter-
mine the applicability of regulatory requirements prior to use. AC 23 Powerplant Guide for Certification of Part 23 Air-
1.5 This international standard was developed in accor- planes and Airships
dance with internationally recognized principles on standard- AC 23.1521-2 Type Certification of Oxygenates and Oxy-
ization established in the Decision on Principles for the genated Gasoline Fuels in Part 23 Airplanes with Recip-
rocating Engines
Development of International Standards, Guides and Recom-
mendations issued by the World Trade Organization Technical
2.3 EASA Requirements:
Barriers to Trade (TBT) Committee.
CS-22 Sailplanes and Powered Sailplanes
CS-E Engines
2. Referenced Documents
CS-P Propellers
2.1 ASTM Standards:
2.4 Other Standards:
D910 Specification for Leaded Aviation Gasolines
EN 228 Automotive Fuels - Unleaded Petrol - Requirements
D4814 Specification for Automotive Spark-Ignition Engine
and Test Methods
Fuel
GAMA Specification No. 1 Specification for Pilot’s Operat-
D7547 Specification for Hydrocarbon Unleaded Aviation
ing Handbook
Gasoline
This specification is under the jurisdiction of ASTM Committee F37 on Light
Sport Aircraft and is the direct responsibility of Subcommittee F37.20 on Airplane. Available from Federal Aviation Administration (FAA), 800 Independence
Current edition approved July 1, 2023. Published August 2023. Originally Ave., SW, Washington, DC 20591, http://www.faa.gov or http://www.gpo.gov.
approved in 2004. Last previous edition approved in 2020 as F2245 – 20. DOI: Available from EASA European Aviation Safety Agency, Postfach 10 12 53,
10.1520/F2245-23. D-50452 Cologne, Germany, http://easa.europa.eu.
2 5
For referenced ASTM standards, visit the ASTM website, www.astm.org, or Available from Deutsches Institut für Normung e.V.(DIN), Am DIN-Platz,
contact ASTM Customer Service at service@astm.org. For Annual Book of ASTM Burggrafenstrasse 6, 10787 Berlin, Germany, http://www.din.de.
Standards volume information, refer to the standard’s Document Summary page on Available from the General Aviation Manufacturers Association (GAMA),
the ASTM website. 1400 K Street NW, Suite 801, Washington, DC 20005-2485, http://www.gama.aero/.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F2245 − 23
3. Terminology 3.2.12 IAS—indicated air speed (m/s, kts)
3.2.13 ICAO—International Civil Aviation Organization
3.1 Definitions:
3.1.1 electric propulsion unit, EPU—any electric motor and
3.2.14 LSA—Light Sport Aircraft
all associated devices used to provide thrust for an electric
3.2.15 MAC—mean aerodynamic chord (m)
aircraft.
3.2.16 n—load factor
3.1.2 energy storage device, ESD—used to store energy as
3.2.17 n —airplane positive maneuvering limit load factor
part of a Electric Propulsion Unit (EPU). Typical energy
storage devices include but are not limited to batteries, fuel
3.2.18 n —airplane negative maneuvering limit load factor
cells, or capacitors.
3.2.19 n —load factor on wheels
3.1.3 flaps—any movable high lift device.
3.2.20 P—power, (kW)
3.1.4 maximum empty weight, W (N)—largest empty 3
E
3.2.21 ρ—air density (kg/m ) = 1.225 at sea level standard
weight of the airplane, including all operational equipment that
conditions
is installed in the airplane: weight of the airframe, powerplant,
3.2.22 POH—Pilot Operating Handbook
Energy Storage Device (ESD) as part of an Electric Propulsion
Unit (EPU), required equipment, optional and specific
2 2
3.2.23 q—dynamic pressure ~N/m !5 ρV
equipment, fixed ballast, full engine coolant and oil, hydraulic
3.2.24 RC—climb rate (m/s)
fluid, and the unusable fuel. Hence, the maximum empty
weight equals maximum takeoff weight minus minimum useful
3.2.25 S—wing area (m )
load: W = W − W .
E U
3.2.26 V—airspeed (m/s)
3.1.5 minimum useful load, W (N)—where W = W − W .
U U E
3.2.26.1 V —design maneuvering speed
A
3.1.6 night—hours between the end of evening civil twilight
3.2.26.2 V —design cruising speed
C
and the beginning of morning civil twilight.
3.2.26.3 V —design diving speed
D
3.1.6.1 Discussion—Civil twilight ends in the evening when
the center of the sun’s disc is 6° below the horizon, and begins 3.2.26.4 V —demonstrated flight diving speed
DF
in the morning when the center of the sun’s disc is 6° below the
3.2.26.5 V —design flap speed
F
horizon.
3.2.26.6 V —maximum flap extended speed
FE
3.1.7 The terms “engine,” referring to internal combustion
3.2.26.7 V —maximum speed in level flight with maximum
H
engines, and “motor,” referring to electric motors for
continuous power (corrected for sea level standard conditions)
propulsion, are used interchangeably within this specification.
3.2.26.8 V —never exceed speed
NE
3.1.8 The term “engine idle,” when in reference to electric
propulsion units, shall mean the minimum power or propeller 3.2.26.9 V —operating maneuvering speed
O
rotational speed condition for the electric motor, as defined
3.2.26.10 V —stalling speed or minimum steady flight
S
without electronic braking of the propeller rotational speed.
speed at which the airplane is controllable (flaps retracted)
3.1.9 The term “vapor lock,” when used in reference to
3.2.26.11 V —stalling speed or minimum steady flight
S1
liquid fuel systems, shall mean that the liquid fuel, while still
speed at which the aircraft is controllable in a specific
in the fuel delivery system, changes state from liquid to gas
configuration
(that is, vaporizes), that either causes fuel feed pressure to the
3.2.26.12 V —stalling speed or minimum steady flight
propulsion unit to decrease below manufacturers’ S0
speed at which the aircraft is controllable in the landing
specifications, transient loss of power, or complete stalling of
configuration
the propulsion unit.
3.2 Abbreviations:
3.2.26.13 V —ground gust speed
R
b
3.2.1 AR—aspect ratio 5 3.2.26.14 V —speed for best angle of climb
X
S
3.2.26.15 V —speed for best rate of climb
3.2.2 b—wing span (m) Y
3.2.27 w—average design surface load (N/m )
3.2.3 c—chord (m)
3.2.28 W—maximum takeoff or maximum design weight
3.2.4 CAS—calibrated air speed (m/s, kts)
(N)
3.2.5 C —lift coefficient of the airplane
L
3.2.29 W —maximum empty airplane weight (N)
E
3.2.6 C —drag coefficient of the airplane
D
3.2.30 W —minimum useful load (N)
U
3.2.7 CG—center of gravity
3.2.31 W —maximum zero wing fuel weight (N)
ZWF
3.2.8 C —moment coefficient (C is with respect to c/4
m m
point, positive nose up)
4. Flight
3.2.9 C —zero lift moment coefficient
MO
4.1 Proof of Compliance:
3.2.10 C —normal coefficient
n
4.1.1 Each of the following requirements shall be met at the
3.2.11 g—acceleration as a result of gravity = 9.81 m/s most critical weight and CG configuration. Unless otherwise
F2245 − 23
TABLE 1 Pilot Force
specified, the speed range from stall to V or the maximum
DF
allowable speed for the configuration being investigated shall Pitch, Roll, Yaw,
Pilot force as applied to the controls
N (lbf) N (lbf) N (lbf)
be considered.
For temporary application (less than 2 min):
4.1.1.1 V may be less than or equal to V .
DF D
Stick 200 (45) 100 (22.5) {
4.1.1.2 V must be less than or equal to 0.9V and greater
NE DF Wheel (applied to rim) 200 (45) 100 (22.5) {
Rudder pedal { { 400 (90)
than or equal to 1.1V . In addition, V must be greater than or
C NE
For prolonged application: 23 (5.2) 23 (5.2) 110 (24.7)
equal to V .
H
4.1.2 The following tolerances are acceptable during flight
testing:
4.4.4 Landing—For landing with throttle closed and flaps
Weight +5 %, −10 %
extended, the following shall be determined:
Weight, when critical +5 %, −1 %
4.4.4.1 Landing distance from 15 m (50 ft) above ground
CG ±7 % of total travel
when speed at 15 m (50 ft) is 1.3V .
SO
4.2 Load Distribution Limits:
4.4.4.2 Ground roll distance with reasonable braking if so
4.2.1 The minimum useful load, W , shall be equal to or
U
equipped.
greater than the sum of:
4.4.5 Balked Landing—The airplane shall demonstrate a
4.2.1.1 An occupant weight of 845 N (190 lbf) for each
full-throttle climb gradient at 1.3 V which shall exceed ⁄30
SO
occupant seat in aircraft, plus
within 5 s of power application from aborted landing. If the
4.2.1.2 The weight of consumable substances, such as fuel,
flaps may be promptly and safely retracted without loss of
as required for a 1 h flight at V . Consumption rates must be
h
altitude and without sudden changes in attitude, they may be
based on test results for the specific application.
retracted.
4.2.2 The minimum flying weight shall be determined.
4.4.5.1 Airplanes with EPU—Balked landing performance
4.2.3 Empty CG, most forward, and most rearward CG shall
shall be demonstrated considering minimum remaining avail-
be determined.
able ESD power.
4.2.4 Fixed or removable ballast, or both, may be used if
4.5 Controllability and Maneuverability:
properly installed and placarded.
4.5.1 General:
4.2.5 Multiple ESDs may be used if properly installed and
4.5.1.1 The airplane shall be safely controllable and maneu-
placarded.
verable during takeoff, climb, level flight (cruise), dive to V
DF
4.3 Propeller Speed and Pitch Limits—Propeller configura-
or the maximum allowable speed for the configuration being
tion shall not allow the engine to exceed safe operating limits
investigated, approach, and landing (power off and on, flaps
established by the engine manufacturer under normal condi-
retracted and extended) through the normal use of primary
tions.
controls.
4.3.1 Maximum RPM shall not be exceeded with full
4.5.1.2 Smooth transition between all flight conditions shall
throttle during takeoff, climb, or flight at 0.9V , and 110 %
H be possible without exceeding pilot force as shown in Table 1.
maximum continuous RPM shall not be exceeded during a
4.5.1.3 Full control shall be maintained when retracting and
glide at V with throttle closed.
NE extending flaps within their normal operating speed range (V
SO
to V ).
4.4 Performance, General—All performance requirements FE
4.5.1.4 Lateral, directional, and longitudinal control shall be
apply in standard ICAO atmosphere in still air conditions and
possible down to V .
at sea level. Speeds shall be given in indicated (IAS) and SO
4.5.2 Longitudinal Control:
calibrated (CAS) airspeeds.
4.5.2.1 With the airplane trimmed as closely as possible for
4.4.1 Stalling Speeds—Wing level stalling speeds V and
SO
steady flight at 1.3V , it must be possible at any speed between
S1
V shall be determined by flight test at a rate of speed decrease
S
1.1V and 1.3V to pitch the nose downward so that a speed
of 0.5 m/s (m/s per second) (1 kt/s) or less, throttle closed, with S1 S1
not less than 1.3V can be reached promptly. This must be
S1
maximum takeoff weight, and most unfavorable CG.
shown with the airplane in all possible configurations, with
4.4.2 Takeoff—With the airplane at maximum takeoff
simultaneous application of full power and nose down pitch
weight, full throttle, the following shall be measured using
control, and with power at idle.
normal takeoff procedures:
4.5.2.2 Longitudinal control forces shall increase with in-
NOTE 1—The procedure used for normal takeoff, including flap
creasing load factor.
position, shall be specified within the POH.
4.5.2.3 The control force to achieve the positive limit
4.4.2.1 Ground roll distance to takeoff on a runway with
maneuvering load factor (n ) shall not be less than 70 N in the
minimal grade.
clean configuration at the aft center of gravity limit. The
4.4.2.2 Distance to clear a 15 m (50 ft) obstacle at a climb
control force increase is to be measured in flight from an initial
speed of at least 1.3V .
n=1 trimmed flight condition at a minimum airspeed of two
S1
4.4.3 Climb—At maximum takeoff weight, flaps in the
times the calibrated maximum flaps up stall speed.
position specified for climb within the POH, and full throttle:
4.5.2.4 If flight tests are unable to demonstrate a maneuver-
4.4.3.1 Rate of climb at V shall exceed 1.6 m/s (315 ing load factor of n , then the minimum control force shall be
Y 1
ft/min).
proportional to the maximum demonstrated load factor, n , as
1D
4.4.3.2 Climb gradient at V shall exceed ⁄12 . follows:
X
F2245 − 23
n 2 1 4.5.5.5 The airplane shall demonstrate compliance with this
1D
f $ 70N
S D
min
n 2 1 section while in trimmed steady flight for each flap and power
setting appropriate to the following configurations: (1) climb
4.5.3 Directional and Lateral Control:
(flaps as appropriate and maximum continuous power); (2)
4.5.3.1 It must be possible to reverse a steady 30° banked
cruise (flaps retracted and 75 % maximum continuous power);
coordinated turn through an angle of 60°, from both directions:
and (3) approach to landing (flaps fully extended and engine at
(1) within 5 s from initiation of roll reversal, with the airplane
idle).
trimmed as closely as possible to 1.3 V , flaps in the takeoff
S1
4.5.6 Dynamic Stability—Any oscillations shall exhibit de-
position, and maximum takeoff power; and (2) within 4 s from
creasing amplitude within the appropriate speed range (1.1 V
S1
initiation of roll reversal, with the airplane trimmed as closely
to maximum allowable speed specified in the POH, both as
as possible to 1.3 V , flaps fully extended, and engine at idle.
SO
appropriate to the configuration).
4.5.3.2 With and without flaps deployed, rapid entry into, or
4.5.7 Wings Level Stall—It shall be possible to prevent more
recovery from, a maximum cross-controlled slip shall not
than 20° of roll or yaw by normal use of the controls during the
result in uncontrollable flight characteristics.
stall and the recovery at all weight and CG combinations.
4.5.3.3 Lateral and directional control forces shall not re-
4.5.8 Turning Flight and Accelerated Turning Stalls:
verse with increased deflection.
4.5.8.1 With the airplane initially trimmed for 1.5 V ,
S
4.5.4 Static Longitudinal Stability:
turning flight and accelerated turning stalls shall be performed
4.5.4.1 The airplane shall demonstrate the ability to trim for
in both directions as follows: While maintaining a 30° coordi-
steady flight at speeds appropriate to the climb, cruise, and
nated turn, apply sufficient pitch control to maintain the
landing approach configurations; at minimum and maximum
required rate of speed reduction until the stall is achieved. After
weight; and forward and aft CG limits.
the stall, level flight shall be regained without exceeding 60° of
4.5.4.2 The airplane shall exhibit positive longitudinal sta-
additional roll in either direction. No excessive loss of altitude
bility characteristics at any speed above 1.1 V , up to the
nor tendency to spin shall be associated with the recovery. The
S1
maximum allowable speed for the configuration being
rate of speed reduction must be nearly constant and shall not
investigated, and at the most critical power setting and CG
exceed 0.5 m/s (m/s per second) (1 kt/s) for turning flight
combination.
stalls and shall be 1.5 to 2.5 m/s (m/s per second) (3 to 5 kt/s)
4.5.4.3 Stability shall be shown by a tendency for the for accelerated turning stalls. The rate of speed reduction in
airplane to return toward trimmed steady flight after: (1) a both cases is controlled by the pitch control.
“push” from trimmed flight that results in a speed increase, 4.5.8.2 Both turning flight and accelerated turning stalls
followed by a non-abrupt release of the pitch control; and (2)
shall be performed: (1) with flaps retracted, at 75 % maximum
a “pull” from trimmed flight that results in a speed decrease, continuous power and at idle; and (2) with flaps extended, at
followed by a non-abrupt release of the pitch control.
75 % maximum continuous power and at idle (speed not to
exceed V ).
4.5.4.4 The airplane shall demonstrate compliance with this
FE
section while in trimmed steady flight for each flap and power (1) Flaps extended conditions include fully extended and
each intermediate normal operating position.
setting appropriate to the following configurations: (1) climb
(flaps set as appropriate and maximum continuous power); (2) (2) If 75 % of maximum continuous power results in pitch
attitudes greater than 30° for non-aerobatic aircraft, the power
cruise (flaps retracted and 75 % maximum continuous power);
and (3) approach to landing (flaps fully extended and engine at setting may be reduced as necessary as follows, but in no case
be less than 50 % of maximum continuous power.
idle).
(a) For flaps retracted, the power setting may be reduced
4.5.4.5 While returning toward trimmed steady flight, the
as necessary to not exceed 30° pitch attitude.
airplane shall: (1) not decelerate below stalling speed V ; (2)
S1
(b) For any flap extended condition, the test may be
not exceed V or the maximum allowable speed for the
NE
carried out with the power required for level flight in the
configuration being investigated; and (3) exhibit decreasing
respective configuration at maximum landing weight and a
amplitude for any long-period oscillations.
speed of 1.4 Vs1.
4.5.5 Static Directional and Lateral Stability:
4.5.5.1 The airplane must maintain a trimmed condition
NOTE 2—If the power setting was reduced to prevent exceeding 30°
pitch attitude, then the POH or Flight Training Supplement must note that
around the roll and yaw axis with respective controls fixed.
the aircraft is not approved for pitch attitudes greater than 30°.
4.5.5.2 The airplane shall exhibit positive directional and
lateral stability characteristics at any speed above 1.2 V , up to 4.5.9 Spinning:
S1
the maximum allowable speed for the configuration being
4.5.9.1 For airplanes placarded “no intentional spins,” the
investigated, and at the most critical power setting and CG
airplane must be able to recover from a one-turn spin or a 3-s
combination.
spin, whichever takes longer, in not more than one additional
turn, with the controls used in the manner normally used for
4.5.5.3 Directional stability shall be shown by a tendency
for the airplane to recover from a skid condition after release of recovery.
the yaw control. 4.5.9.2 For airplanes in which intentional spinning is
allowed, the airplane must be able to recover from a three-turn
4.5.5.4 Lateral stability shall be shown by a tendency for the
spin in not more than one and one-half additional turn.
airplane to return toward a level-wing attitude after release of
the roll control from a slip condition. 4.5.9.3 In addition, for either 4.5.9.1 or 4.5.9.2:
F2245 − 23
(1) For both the flaps-retracted and flaps-extended 5.1.2.1 Unless otherwise provided in 5.1.2.2, an ultimate
conditions, the applicable airspeed limit and limit maneuvering load factor of safety of 1.5 must be used.
load factor may not be exceeded.
5.1.2.2 Special ultimate load factors of safety shall be
(2) There may be no excessive control forces during the
applied to the following:
spin or recovery.
2.0 × 1.5 = 3.0 on castings
(3) It must be impossible to obtain uncontrollable spins 1.2 × 1.5 = 1.8 on fittings
2.0 × 1.5 = 3.0 on bearings at bolted or pinned joints subject to rotation
with any use of the controls.
4.45 × 1.5 = 6.67 on control surface hinge-bearing loads except ball
(4) For the flaps-extended condition, the flaps may be
and roller bearing hinges
retracted during recovery. 2.2 × 1.5 = 3.3 on push-pull control system joints
1.33 × 1.5 = 2 on cable control system joints, lap belt/shoulder harness fittings
4.5.9.4 For those airplanes of which the design is inherently
(including the seat if belt/harness is attached to it)
spin resistant, such resistance must be proven by test and
5.1.3 Strength and Deformation:
documented. If proven spin resistant, the airplane must be
5.1.3.1 The structure must be able to support limit loads
placarded “no intentional spins” but need not comply with
without detrimental, permanent deformation. At any load up to
4.5.9.1 – 4.5.9.3.
limit loads, the deformation shall not interfere with safe
4.6 Vibrations—For airplanes with V equal to or less than
H
operation.
120 (CAS), flight testing shall not reveal, by pilot observation,
5.1.3.2 The structure must be able to support ultimate loads
heavy buffeting (except as associated with a stall), excessive
without failure for at least 3 s. However, when proof of
airframe or control vibrations, flutter (with proper attempts to
strength is shown by dynamic tests simulating actual load
induce it), or control divergence, at any speed from V to V .
SO DF
conditions, the 3-s limit does not apply.
If V is greater than 120 (CAS), use Specification F3619,
H
5.1.4 Proof of Structure—Each design requirement must be
Flutter Standard.
verified by means of conservative analysis or test (static,
4.7 Ground and Water Control and Stability:
component, or flight), or both.
4.7.1 It must be possible to taxi, takeoff, and land while
5.1.4.1 Compliance with the strength and deformation re-
maintaining control of the airplane, up to the maximum
quirements of 5.1.3 must be shown for each critical load
crosswind component specified within the POH.
condition. Structural analysis may be used only if the structure
4.7.2 Wheel brakes must operate so as not to cause unpre-
conforms to those for which experience has shown this method
dictable airplane response or control difficulties.
to be reliable. In other cases, substantiating load tests must be
4.7.3 A seaplane or amphibian may not have dangerous or
made. Dynamic tests, including structural flight tests, are
uncontrollable porpoising characteristics at any normal oper-
acceptable if the design load conditions have been simulated.
ating speed on the water.
Substantiating load tests should normally be taken to ultimate
4.8 Spray Characteristics—Spray may not dangerously ob-
design load.
scure the vision of the pilots or damage the propeller or other
5.1.4.2 Certain parts of the structure must be tested as
critical parts of a seaplane or amphibian at any time during
specified in 6.9.
taxiing, take-off, and landing.
5.2 Flight Loads:
5.2.1 General:
5. Structure
5.2.1.1 Flight load factors, n, represent the ratio of the
5.1 General:
aerodynamic force component (acting normal to the assumed
5.1.1 Loads:
longitudinal axis of the airplane) to the weight of the airplane.
5.1.1.1 Strength requirements are specified in terms of limit
A positive flight load factor is one in which the aerodynamic
loads (the maximum loads to be expected in service) and
force acts upward, with respect to the airplane.
ultimate loads (limit loads multiplied by prescribed factors of
5.2.1.2 Compliance with the flight load requirements of this
safety). Unless otherwise provided, prescribed loads are limit
section must be shown at each critical weight distribution
loads.
within the operating limitations specified in the POH.
5.1.1.2 Unless otherwise provided, the air, ground, and
5.2.1.3 Maximum Zero Wing Fuel Weight, W —The
ZWF
water loads must be placed in equilibrium with inertia forces,
maximum allowable weight of the airplane without any fuel in
considering each item of mass in the airplane. These loads must
the wing tank(s) must be established if it is less than maximum
be distributed to conservatively approximate or closely repre-
design weight, W.
sent actual conditions.
5.2.2 Symmetrical Flight Conditions:
5.1.1.3 If deflections under load would significantly change
5.2.2.1 The appropriate balancing horizontal tail loads must
the distribution of external or internal loads, this redistribution
must be taken into account. be accounted for in a rational or conservative manner when
determining the wing loads and linear inertia loads correspond-
5.1.1.4 Appendix X1 – Appendix X5 provide, within the
ing to any of the symmetrical flight conditions specified in
limitations specified within the appendix, a simplified means of
5.2.2 to 5.2.6.
compliance with several of the requirements set forth in 5.2.1
to 5.7.3 that can be applied as one (but not the only) means to 5.2.2.2 The incremental horizontal tail loads due to maneu-
comply.
vering and gusts must be reacted by the angular inertia of the
5.1.2 Factor of Safety: airplane in a rational or conservative manner.
F2245 − 23
5.2.2.3 In computing the loads arising in the conditions
V 5 V ·=n
A S 1
prescribed above, the angle of attack is assumed to be changed
W
suddenly without loss of air speed until the prescribed load
V 5 , m/s
~ !
S
factor is attained. Angular accelerations may be disregarded.
ρC S
!
LMAX
5.2.2.4 The aerodynamic data required for establishing the
loading conditions must be verified by tests, calculations, or by
where:
conservative estimation. In the absence of better information,
V = computed stalling speed at the design maximum weight
S
the maximum negative lift coefficient for rigid lifting surfaces
with the flaps retracted, and
may be assumed to be equal to −0.80. If the pitching moment
n = positive limit maneuvering load factor used in design.
coefficient, C , is less than 60.025, a coefficient of at least
mo
5.2.4.2 Design Flap Speed, V —For each landing setting,
60.025 must be used.
F
V must not be less than the greater of: (1) 1.4 V , where V is
5.2.3 Flight Loads Envelope (V-n Diagram)—Compliance F S S
the computed stalling speed with the wing flaps retracted at the
shall be shown at any combination of airspeed and load factor
maximum weight; and (2) 2.0 V , where V is the computed
on the boundaries of the flight loads envelope. The flight loads SO SO
stalling speed with wing flaps fully extended at the maximum
envelope represents the envelope of the flight loading condi-
weight.
tions specified by the criteria of 5.2.4 and 5.2.5 (see Fig. 1).
5.2.4.3 Design Cruising Speed, V —(1) V may not be less
5.2.3.1 General—Compliance with the strength require-
C C
ments of this subpart must be shown at any combination of
than 2.45=W/S ; and (2) V need not be greater than 0.9 V at
C H
airspeed and load factor on and within the boundaries of a sea level.
flight loads envelope similar to the one in Fig. 1 that represents
5.2.4.4 Design Dive Speed, V :
D
the envelope of the flight loading conditions specified by the
V 5 1.4 × V
maneuvering and gust criteria of 5.2.5 and 5.2.6 respectively. D Cmin
5.2.3.2 Maneuvering Envelope—Except where limited by
where:
maximum (static) lift coefficients, the airplane is assumed to be
V = required minimum cruising speed.
C min
subjected to symmetrical maneuvers resulting in the following
5.2.5 Limit Maneuvering Load Factors:
limit load factors: (1) the positive maneuvering load factor
5.2.5.1 The positive limit maneuvering load factor n may
specified in 5.2.5.1 at speeds up to V ; and (2) the negative 1
D
not be less than 4.0.
maneuvering load factor specified in 5.2.5.2 at speeds up to V .
D
5.2.5.2 The negative limit maneuvering load factor n may
5.2.3.3 Gust Envelope—The airplane is assumed to be
not be greater than −2.0.
subjected to symmetrical vertical gusts in level flight. The
5.2.5.3 Loads with wing flaps extended: (1) if flaps or other
resulting limit load factors must correspond to the conditions
similar high lift devices are used, the airplane must be designed
determined as follows: (1) positive (up) and negative (down)
for n = 2.0 with the flaps in any position up to V ; and (2) n
gusts of 15 m/s (49.2 ft/s) at V ; and (2) positive and negative
1 F 2
C
= 0.
gusts of 7.5 m/s (24.6 ft/s) at V (see Fig. 1).
D
5.2.4 Design Airspeeds: 5.2.5.4 Loads with speed control devices: (1) if speed
5.2.4.1 Design Maneuvering Speed, V : control devices such as speed brakes or spoilers are used, the
A
FIG. 1 Flight Loads Envelope (V-n Diagram)
F2245 − 23
airplane must be designed for a positive limit load factor of 3.0 (2) 2, 3, 4, or 8 for engines with four, three, two, or one
with the devices extended in any position up to the placard cylinders, respectively.
device extended speed; and (2) maneuvering load factors lower
For two-stroke engines:
than those specified in 5.2.5 may be used if the airplane has
(1) 2 for engines with three or more cylinders; or
design features that make it impossible to exceed these in
(2) 3 or 6, for engines with two or one cylinders, respec-
flight.
tively.
5.2.6 Gust Load Factors—The airplane must be designed
5.2.9.4 For conventional electric motors with positive drive
for the loads resulting from:
to the propeller, the limit torque to be accounted for in 5.2.9.1
5.2.6.1 The gust velocities specified in 5.2.3.3 with flaps
and 5.2.9.2 is obtained by multiplying the mean torque by 1.33.
retracted, and
5.2.10 Side Load on Engine Mount:
5.2.6.2 Positive and negative gusts of 7.5 m/s (24.6 ft/s)
5.2.10.1 The engine mount and its supporting structure must
nominal intensity at V with the flaps fully extended.
F
be designed for a limit load factor in a lateral direction, for the
side load on the engine mount, of not less than 1.5.
NOTE 3—In the absence of a more rational analysis, the gust load
factors may be computed by the method of Appendix X4. 5.2.10.2 The side load prescribed in 5.2.10.1 may be as-
sumed to be independent other flight conditions.
5.2.7 Unsymmetrical Flight Conditions—The airplane is
5.2.10.3 If applicable, the nose wheel loads of 5.8.1.7 must
assumed to be subjected to the unsymmetrical flight conditions
also be considered.
of 5.2.7.1 and 5.2.7.2. Unbalanced aerodynamic moments
about the center of gravity must be reacted in a rational or
5.3 Control Surface and System Loads:
conservative manner considering the principal masses furnish-
5.3.1 Control Surface Loads—The control surface loads
ing the reacting inertia forces.
specified in 5.3.3 through 5.7.3 are assumed to occur in the
5.2.7.1 Rolling Conditions—The airplane shall be designed
conditions described in 5.2.2 through 5.2.6.
for the loads resulting from the roll control deflections and
5.3.2 Control System Loads—Each part of the primary
speeds specified in 5.7.1 in combination with a load factor of
control system situated between the stops and the control
at least two thirds of the positive maneuvering load factor
surfaces must be designed for the loads corresponding to at
prescribed in 5.2.5.1. The rolling accelerations may be ob-
least 125 % of the computed hinge moments of the movable
tained by the methods given in X3.4. The effect of the roll
control surfaces resulting from the loads in the conditions
control displacement on the wing torsion may be accounted for
prescribed in 5.3.1 through 5.7.3. In computing the hinge
by the method of X3.4.2 and X3.4.3.
moments, reliable aerodynamic data must be used. In no case
5.2.7.2 Yawing Conditions—The airplane must be designed
may the load in any part of the system be less than those
for the yawing loads resulting from the vertical surface loads
resulting from the application of 60 % of the pilot forces
specified in 5.5.
described in 5.3.3. In addition, the system limit loads need not
5.2.8 Special Conditions for Rear Lift Truss:
exceed the loads that can be produced by the pilot. Pilot forces
5.2.8.1 If a rear lift truss is used, it must be designed for
used for design need not exceed the maximum pilot forces
conditions of reversed air flow at a design speed of:
prescribed in 5.3.3.
5.3.3 Loads Resulting from Limit Pilot Forces—The main
W
V 5 0.65Π14.5
control systems for the direct control of the airplane about its
S
longitudinal, lateral, or yaw axis, including the supporting
where:
points and stops, must be designed for the limit loads resulting
W/S = wing loading, N/m . from the limit pilot forces as follows:
5.3.3.1 Pitch—445 N (100 lbf) at the grips of the stick or
5.2.8.2 Either aerodynamic data for the particular wing
wheel.
section used, or a value of C equaling −0.8 with a chord-wise
L
5.3.3.2 Roll—180 N (40.5 lbf) at the grip(s) of the stick or
distribution that is triangular between a peak at the trailing
wheel.
edge and zero at the leading edge, must be used.
5.3.3.3 Yaw—580 N (130 lbf) acting forward on one rudder
5.2.9 Engine Torque—The engine mount and its supporting
pedal.
structure must be designed for the effects of:
5.3.3.4 The rudder control system must be designed to a
5.2.9.1 The limit torque corresponding to takeoff power and
load of 580 N (130 lbf) per pedal acting simultaneously on both
propeller speed acting simultaneously with 75 % of the limit
pedals in the forward direction.
loads from flight condition of 5.2.5.1.
5.3.4 Dual-Control Systems—Dual-control systems must be
5.2.9.2 The limit torque corresponding to maximum con-
designed for the loads resulting from each pilot applying 0.75
tinuous power and propeller speed acting simultaneously with
times the load specified in 5.3.3 with the pilots acting in
the limit loads from flight condition of 5.2.5.1.
opposition.
5.2.9.3 For conventional reciprocating engines with positive
5.3.5 Secondary Control Systems—Secondary control
drive to the propeller, the limit torque to be accounted for in
5.2.9.1 and 5.2.9.2 is obtained by multiplying the mean torque systems, such as those for flaps and trim control must be
designed for the maximum forces that a pilot is likely to apply.
by one of the following factors:
For four-stroke engines: 5.3.6 Control System Stiffness and Stretch—The amount of
(1) 1.33 for engines with five or more cylinders; or control surface or tab movement available to the pilot shall not
F2245 − 23
NOTE 5—In the absence of a more rational analysis, the horizontal
be dangerously reduced by elastic stretch or shortening of the
surfaces gust loads may be computed by the method of Appendix X5.
system in any condition.
5.3.7 Ground Gust Conditions—The control system from
5.5 Vertical Stabilizing Surfaces:
the control surfaces to the stops or control locks, when
5.5.1 Maneuvering Loads—The vertical stabilizing surfaces
installed, must be designed for limit loads due to gusts
must be designed for maneuvering loads imposed by the
corresponding to the following hinge moments:
following conditions:
5.5.1.1 At speed V , full deflection of the yaw control in
M 5 k·C ·S ·q (1) A
S S S
both directions.
where:
5.5.1.2 At speed V , one-third full deflection of the yaw
D
M = limit hinge moment,
S control in both directions.
C = mean chord of the control surface aft of the hinge line,
S
5.5.2 Gust Loads:
S = area of the control surface aft of the hinge line,
S
5.5.2.1 The vertical stabilizing surfaces must be designed to
q = dynamic pressure corresponding to an airspeed of 20
withstand lateral gusts of the values prescribed in 5.2.3.3.
m/s (38 kts), and
k = limit hinge moment coefficient due to ground gust =
NOTE 6—In the absence of a more rational analysis, the vertical surfaces
gust loads may be computed by the method in Appendix X5.2.
0.75.
5.5.3 Outboard Fins or Winglets:
5.3.8 Control Surface Mass Balance Weights—If applicable
shall be designed for: 5.5.3.1 If outboard fins or winglets are on the horizontal
surfaces or wings, the horizontal surfaces or wings must be
5.3.8.1 The n = 16 limit load normal to the surface, and
5.3.8.2 The n = 8 limit load fore and aft and parallel to the designed for their maximum load in combination with loads
induced by the fins or winglets and moments or forces exerted
hinge line.
on the horizontal surfaces or wings by the fins or winglets.
5.3.9 The motion of wing flaps on opposite sides of the
plane of symmetry must be synchronized by a mechanical 5.5.3.2 If outboard fins or winglets extend above and below
interconnection unless the airplane has safe flight characteris- the horizontal surface, the critical vertical surface loading (the
tics with the wing flaps retracted on one side and extended on load per unit area determined in accordance with 5.5.1 and
the other. 5.5.2) must be applied to:
5.3.10 All primary controls shall have stops within the (1) The part of the vertical surface above the horizontal
system to withstand the greater of pilot force, 125 % of surface surface with 80 % of that loading applied to the part below the
horizontal surface or wing, and
loads, or ground gust loads (see 5.3.7).
(2) The part of the vertical surface below the horizontal
5.4 Horizontal Stabilizing and Balancing Surfaces:
surface or wing with 80 % of that loading applied to the part
5.4.1 Balancing Loads:
above the horizontal surface or wing.
5.4.1.1 A horizontal stabilizing surface balancing load is the
5.5.3.3 The end plate effects of outboard fins or winglets
load necessary to maintain equilibrium in any specified flight
must be taken into account in applying the yawing conditions
condition with no pitching acceleration.
of 5.5.1 and 5.5.2 to the vertical surfaces in 5.5.3.2.
5.4.1.2 Horizontal stabilizing surfaces must be designed for
5.5.3.4 When rational methods are used for computing
the balancing loads occurring at any point on the limit
loads, the maneuvering loads of 5.5.1 on the vertical surfaces
maneuvering envelope and in the air-brake and wing-flap
and the n = 1 horizontal surface or wing load, including
positions specified in 5.2.5.3.
induced loads on the horizontal surface or wing and moments
5.4.2 Maneuvering Loads—Horizontal stabilizing surfaces
or forces exerted on the horizontal surfaces or wing, must be
must be designed for pilot-induced pitching maneuvers im-
applied simultaneously for the structural loading condition.
posed by the following conditions:
5.4.2.1 At speed V , maximum upward deflection of pitch
A 5.6 Supplementary Conditions for Stabilizing Surfaces:
control surface,
5.6.1 Combined Loads on Stabilizing Surfaces:
5.4.2.2 At speed V , maximum downward deflection of
A
5.6.1.1 With the airplane in a loading condition correspond-
pitch control surface,
ing to A or D in Fig. 1 (whichever condition leads to the higher
5.4.2.3 At speed V , one-third maximum upward deflection
D
balance load) the loads on the horizontal surface must be
of pitch control surface, and
combined with those on the vertical surface as specified in
5.4.2.4 At speed V , one-third maximum downward deflec-
D
5.5.1. It must be assumed that 75 % of the loads according to
tion of pitch control surface.
5.4.2 for the horizontal stabilizing surface and 5.5.1 for the
vertical stabilizing surface are acting simultaneously.
NOTE 4—In 5.4.2, the following assumptions should be made: the
airplane is initially in level flight, and its altitude and airspeed do not
5.6.1.2 The stabilizing surfaces and fuselage must be de-
change. The loads are balanced by inertia forces.
signed for asymmetric loads on the stabilizing surfaces which
would result from application of the highest symmetric ma-
5.4.3 Gust Loads—The horizontal stabilizing surfaces must
be designed for the loads resulting from: neuver loads of 5.4.2 so that 100 % of the horizontal stabilizer
surface loading is applied to one side of the plane symmetry
5.4.3.1 The gust velocities specified in 5.2.3.3 with flaps
and 70 % on the opposite side.
retracted, and
5.4.3.2 Positive and negative gusts of 7.5 m/s (24.6 ft/s) 5.6.2 Additional Loads Applying to V-Tails—An airplane
nominal intensity at V with the flaps fully extended. with a V-tail must be designed for a gust acting perpendicular
F
F2245 − 23
to one of the surfaces at speed V . This condition is supple- where:
C
mental to the equivalent horizontal and vertical cases previ-
W
h =
drop height, m50.0132 =W⁄S with ⁄S in N/m , but
ously specified.
h larger than 0.23 m (9.1 in.),
5.7 Ailerons, Wing Flaps, and Special Devices:
d = total shock absorber travel, m = d + d ,
tire shock
5.7.1 Ailerons—The ailerons must be designed for control ef = shock efficiency, and
ef × d = 0.5 × d for tire and rubber or spring shocks, or
loads corresponding to the following conditions:
= 0.5 × d + 0.65 × d for hydraulic shock
5.7.1.1 At speed V , the full deflection of the roll control.
tire shock
A
absorbers.
5.7.1.2 At speed V , one-third of the full deflection of the
D
roll control.
5.8.1.2 If n is larger than 3.33, all concentrated masses
j
5.7.2 Flaps—Wing flaps, their operating mechanisms, and
(engine, fuel tanks, occupant seats, ballast, etc.) must be
supporting structure must be designed for the critical loads
substantiated for a limit landing load factor of n + 0.67 = n
j
occurring in the flaps-extended operating range with the flaps
which is greater than 4.
in any position. The effects of propeller slipstream, correspond-
5.8.1.3 The usual ultimate factor of safety of 1.5 applies to
ing to takeoff power, must be taken into account at an airspeed
these conditions, unless a drop test from the reserve energy
of not less than 1.4 V , where V is the computed stalling speed
S S
height, hr = 1.44h, shows that a lower factor may be used. If
with flaps fully retracted at the design weight. For investigating
the shock absorber is of a fast energy absorbing type, the
the slipstream effects, the load factor may be assumed to be
ultimate loads are the limit load multiplied by the conservative
1.0.
reserve energy factor of 1.2.
5.7.3 Special Devices—The loadings for special devices
5.8.1.4 Side Load Conditions—The requirements for the
using aerodynamic surfaces, such as slots and spoilers, must be
side load conditions on the main wheels in a level attitude are
determined from test data or reliable aerodynamic data that
given in Fig. 3.
allows close estimates.
5.8.1.5 Braked Roll Conditions—The requirements for the
5.8 Ground Load Conditions:
braked roll conditions on the main wheels in a level attitude are
5.8.1 Basic Landing Conditions—The requirements for the
given in Fig. 4.
basic landing conditions are given in 5.8.1.1 to 5.8.1.3, Table 2,
5.8.1.6 Supplementary Conditions for Tail Wheel—The re-
and Fig. 2.
quirements for the tail wheel conditions in a tail down attitude
5.8.1.1 The load factor on the wheels, n , may be computed
j
are given in Fig. 5.
as follows:
5.8.1.7 Supplementary Conditions for Nose Wheel—The
d
requirements for supplementary conditions for nose wheels are
h1
3 given in Fig. 6 (the static load is at the combination of weight
n 5
j
and CG that gives the maximum loads).
ef × d
TABLE 2 Basic Landing Conditions
NOTE 1—
K = 0.25
L = ⁄3 = ratio of the assumed w
...