ASTM F2564-14(2022)

Standard Specification for Design and Performance of a Light Sport Glider

ABSTRACT
This specification covers airworthiness requirements for the design of a powered or non-powered fixed wing light sport aircraft, a “glider.” Stability shall be shown by a tendency for the glider to return toward steady flight after: (1) a “push” from steady flight that results in a speed increase, followed by a non-abrupt release of the pitch control; and (2) a “pull” from steady flight that results in a speed decrease, followed by a non-abrupt release of the pitch control. Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). The suitability of each structural design detail and part having an important bearing on safety shall be established by test. Each combination of engine, exhaust, cooling and fuel system on a powered glider must be compatible with the glider, and function in a safe and satisfactory manner within the operational limits of the glider and powerplant. Each aircraft shall include Aircraft Operating Instructions (AOI).
SCOPE
1.1 This specification covers airworthiness requirements for the design of a powered or non-powered fixed wing light sport aircraft, a “glider.”  
1.2 This specification is applicable to the design of a light sport aircraft glider as defined by regulations and limited to day VFR flight.  
1.3 A glider for the purposes of this specification is defined as a heavier than air aircraft that remains airborne through the dynamic reaction of the air with a fixed wing and in which the ability to remain aloft in free flight does not depend on the propulsion from a power plant. A powered glider is defined for the purposes of this specification as a glider equipped with a power plant in which the flight characteristics are those of a glider when the power plant is not in operation.  
1.4 The values stated in SI units are to be regarded as standard. The values given in parenthesis are for information only.  
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 requirements 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.

General Information

• Status: Published
• Publication Date: 30-Sep-2022
• ICS: 49.020 - Aircraft and space vehicles in general
• Technical Committee: F37 - Light Sport Aircraft
• Drafting Committee: F37.10 - Glider

Relations

• Refers: ASTM F2339-19a - Standard Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft
• Effective Date: 01-Nov-2019
• Refers: ASTM F2339-19 - Standard Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft
• Effective Date: 01-Apr-2019
• Refers: ASTM F2339-17 - Standard Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft
• Effective Date: 01-Dec-2017
• Refers: ASTM F2972-15 - Standard Specification for Light Sport Aircraft Manufacturer’s Quality Assurance System (Withdrawn 2024)
• Effective Date: 01-Dec-2015
• Refers: ASTM F2972-14 - Standard Specification for Light Sport Aircraft Manufacturer’s Quality Assurance System
• Effective Date: 01-Nov-2014
• Refers: ASTM F2972-14e1 - Standard Specification for Light Sport Aircraft Manufacturer’s Quality Assurance System
• Effective Date: 01-Nov-2014
• Refers: ASTM F2972-12 - Standard Specification for Light Sport Aircraft Manufacturer’s Quality Assurance System
• Effective Date: 01-Dec-2012
• Refers: ASTM F2316-12 - Standard Specification for Airframe Emergency Parachutes
• Effective Date: 01-Sep-2012
• Refers: ASTM F2840-11 - Standard Practice for Design and Manufacture of Electric Propulsion Units for Light Sport Aircraft
• Effective Date: 01-Mar-2011
• Refers: ASTM F2316-08(2010) - Standard Specification for Airframe Emergency Parachutes for Light Sport Aircraft
• Effective Date: 01-Dec-2010
• Refers: ASTM F2295-10 - Standard Practice for Continued Operational Safety Monitoring of a Light Sport Aircraft (Withdrawn 2019)
• Effective Date: 01-Jan-2010
• Refers: ASTM F2339-06(2009) - Standard Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft
• Effective Date: 01-Jan-2009
• Refers: ASTM F2316-08 - Standard Specification for Airframe Emergency Parachutes for Light Sport Aircraft
• Effective Date: 01-Oct-2008
• Refers: ASTM F2316-06 - Standard Specification for Airframe Emergency Parachutes for Light Sport Aircraft
• Effective Date: 01-Oct-2006
• Refers: ASTM F2295-06 - Standard Practice for Continued Operational Safety Monitoring of a Light Sport Aircraft
• Effective Date: 15-Aug-2006

Overview

ASTM F2564-14(2022) sets forth comprehensive airworthiness and design requirements for powered and non-powered fixed wing light sport aircraft, specifically gliders. Developed by ASTM International, this standard covers stability, strength, and structural requirements to ensure the safe design, performance, and operation of light sport gliders. Applicable to both single and two-seater configurations, ASTM F2564 is intended to foster safety, compatibility, and reliability for manufacturers, operators, and regulators within the sport aviation industry.

Key Topics

• Scope and Definitions: Covers powered and non-powered heavier-than-air glider aircraft with fixed wings. Powered gliders must behave as gliders when the powerplant is off.
• Airworthiness Requirements: Specifies stability criteria, including a glider's ability to return to steady flight after abrupt pitch control movements.
• Strength and Load Factors: Details strength requirements using limit loads (maximum expected in service) and ultimate loads (limit loads with safety factors).
• Structural Suitability: Mandates testing of critical parts and design details for safety.
• Compatibility of Powered Systems: For powered gliders, requires that engine, exhaust, cooling, and fuel systems function safely within operational limits.
• Stability and Control: Establishes requirements for static and dynamic stability in various flight conditions, stall and spin recovery, and control forces.
• Aircraft Operating Instructions (AOI): Every light sport glider must include AOI for safe operation, including procedures for takeoff, climb, flight, descent, and landing.

Applications

ASTM F2564-14(2022) has practical value across multiple aviation stakeholders:

• Manufacturers: Provides a standardized framework for the design and testing of light sport gliders, supporting certification and regulatory compliance.
• Aviation Authorities and Inspectors: Offers clear criteria to assess the airworthiness, safety, and design quality of sport gliders.
• Maintenance Organizations: Establishes structurally critical components and load requirements, aiding in ongoing airworthiness and maintenance planning.
• Operators and Pilots: Ensures that operational instructions reflect tested and approved flight characteristics and limitations, bolstering pilot confidence and safety.
• Education and Training: Serves as a benchmark for curriculum developers and instructors in light sport aviation, ensuring trainees are familiar with current safety and performance standards.

Related Standards

ASTM F2564 is part of a broader suite of standards supporting light sport aircraft. Key related documents include:

• ASTM F2316: Specification for Airframe Emergency Parachutes, relevant for safety systems integration.
• ASTM F2339: Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft.
• ASTM F2840: Practice for Design and Manufacture of Electric Propulsion Units for Light Sport Aircraft.
• ASTM F2972: Specification for Light Sport Aircraft Manufacturer’s Quality Assurance System.
• CS-22 Subpart H: Certification Specifications for Sailplanes and Powered Sailplanes from EASA, providing additional regulatory context.

Conclusion

Adhering to ASTM F2564-14(2022) ensures that light sport gliders meet internationally recognized design and performance criteria. The standard enhances safety, reliability, and interoperability in sport aviation by providing clear requirements for stability, structure, and operational compatibility. Comprehensive implementation benefits designers, regulators, operators, and the entire sports aviation community, promoting trust and safety in recreational and instructional glider flight.

Keywords: ASTM F2564, light sport glider, airworthiness, glider design, sport aviation standards, glider safety, powered glider requirements, structural testing, aircraft stability, aviation compliance

Standards Content (Sample)

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:F2564 −14 (Reapproved 2022)
Standard Specification for
Design and Performance of a Light Sport Glider
This standard is issued under the fixed designation F2564; the number immediately following the designation indicates the year of
original adoption or, in the case of revision, the year of last revision.Anumber in parentheses indicates the year of last reapproval.A
superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope ing of a Light Sport Aircraft (Withdrawn 2019)
F2316Specification for Airframe Emergency Parachutes
1.1 Thisspecificationcoversairworthinessrequirementsfor
F2339Practice for Design and Manufacture of Reciprocat-
the design of a powered or non-powered fixed wing light sport
ing Spark Ignition Engines for Light Sport Aircraft
aircraft, a “glider.”
F2840Practice for Design and Manufacture of Electric
1.2 This specification is applicable to the design of a light
Propulsion Units for Light Sport Aircraft
sportaircraftgliderasdefinedbyregulationsandlimitedtoday
F2972Specification for Light SportAircraft Manufacturer’s
VFR flight.
Quality Assurance System
1.3 Aglider for the purposes of this specification is defined 2.2 Other Standard:
CS-22 Subpart HCertification Specifications for Sailplanes
as a heavier than air aircraft that remains airborne through the
dynamic reaction of the air with a fixed wing and in which the and Powered Sailplanes
ability to remain aloft in free flight does not depend on the
3. Terminology
propulsion from a power plant.Apowered glider is defined for
the purposes of this specification as a glider equipped with a 3.1 Definitions:
power plant in which the flight characteristics are those of a 3.1.1 electric propulsion unit, EPU—any electric motor and
all associated devices used to provide thrust for an electric
glider when the power plant is not in operation.
aircraft.
1.4 The values stated in SI units are to be regarded as
3.1.2 energy storage device, ESD—used to store energy as
standard. The values given in parenthesis are for information
part of a Electric Propulsion Unit (EPU). Typical energy
only.
storage devices include but are not limited to batteries, fuel
1.5 This standard does not purport to address all of the
cells or capacitors.
safety concerns, if any, associated with its use. It is the
3.1.3 feathering—a single action from the cockpit that
responsibility of the user of this standard to establish appro-
repositionsthepropellerbladestolowdragconfigurationwhen
priate safety, health, and environmental practices and deter-
the engine is not operating.
mine the applicability of regulatory requirements prior to use.
1.6 This international standard was developed in accor-
3.1.4 flaps—any movable high lift device.
dance with internationally recognized principles on standard-
3.1.5 maximum empty weight, W (kg)—largest empty
E
ization established in the Decision on Principles for the
weightoftheglider,includingalloperationalequipmentthatis
Development of International Standards, Guides and Recom-
installed in the glider: weight of the airframe, powerplant,
mendations issued by the World Trade Organization Technical
excluding energy storage device (ESD) for electric propulsion
Barriers to Trade (TBT) Committee.
unit when removable, required equipment, optional and spe-
cific equipment, fixed ballast, full engine coolant and oil,
2. Referenced Documents
hydraulic fluid, and the unusable fuel. Hence, the maximum
2.1 ASTM Standards:
emptyweightequalsmaximumtakeoffweightminusminimum
F2295Practice for Continued Operational Safety Monitor-
useful load: W = W– W .
E U
3.1.6 minimum useful load, W (kg)—where W = W – W .
U U E
3.1.7 The terms “engine” referring to internal combustion
This specification is under the jurisdiction ofASTM Committee F37 on Light
enginesand“motor”referringtoelectricmotorsforpropulsion
Sport Aircraft and is the direct responsibility of Subcommittee F37.10 on Glider.
Current edition approved Oct. 1, 2022. Published October 2022. Originally
are used interchangeably within this standard.
approved in 2006. Last previous edition approved in 2014 as F2564–14. DOI:
10.1520/F2564-14R22.
2 3
For referenced ASTM standards, visit the ASTM website, www.astm.org, or The last approved version of this historical standard is referenced on
contact ASTM Customer Service at [e-mail]. For Annual Book of ASTM www.astm.org.
Standards volume information, refer to the standard’s Document Summary page on Available from European Union Aviation Safety Agency (EASA), Konrad-
the ASTM website. Adenauer-Ufer 3, D-50668 Cologne, Germany, https://www.easa.europa.eu.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F2564−14 (2022)
3.1.8 The term “engine idle” or “throttle closed” when in 3.2.35 V —stalling speed or minimum steady flight speed
S0
reference to electric propulsion units shall mean the minimum atwhichtheaircraftiscontrollableinthelandingconfiguration
power or propeller rotational speed condition for the electric
3.2.36 V —ground gust speed
R
motor as defined without electronic braking of the propeller
3.2.37 V —maximum aerotow speed
T
rotational speed.
3.2.38 V —maximum winch tow speed
3.2 Abbreviations:
W
3.2.1 AOI—Aircraft Operating Instructions
3.2.39 V —speed for best rate of climb
Y
3.2.2 AR—Aspect Ratio = b /S
3.2.40 W—maximum takeoff or maximum design weight
3.2.3 b—wing span (m) (kg)
3.2.4 c—chord (m) 3.2.41 W —maximum empty aircraft weight (kg)
E
3.2.5 CAS—calibrated air speed (m/s, kts) 3.2.42 W —minimum useful load (kg)
U
3.2.6 C —lift coefficient of the aircraft 3.2.43 w—average design surface load (N/m )
L
3.2.7 C —drag coefficient of the aircraft
D
4. Flight
3.2.8 CG—center of gravity
4.1 Proof of Compliance:
3.2.9 C —moment coefficient (C is with respect to c/4
m m
4.1.1 Eachofthefollowingrequirementsshallbemetatthe
point, positive nose up)
most critical weight and CG configuration. Unless otherwise
3.2.10 C —zero lift moment coefficient
MO
specified, the speed range from stall to V or the maximum
DF
3.2.11 C —normal coefficient allowable speed for the configuration being investigated shall
n
be considered.
3.2.12 g—acceleration as a result of gravity = 9.81 m/s
4.1.1.1 V shall be less than or equal to V .
DF D
3.2.13 IAS—indicated air speed (m/s, kts)
4.1.1.2 If V chosen is less than V , V must be less than
DF D NE
3.2.14 ICAO—International Civil Aviation Organization
or equal to 0.9 V and greater than or equal to 1.1 V .
DF C
3.2.15 LSA—light sport aircraft 4.1.2 The following tolerances are acceptable during flight
testing:
3.2.16 n—load factor
Weight +5 %, −10 %
3.2.17 n —glider positive maneuvering limit load factor
Weight, when critical +5 %, −1 %
at V
CG ±7 % of total travel
A
3.2.18 n —glider positive maneuvering limit load factor 4.2 Compliance must be established for all configurations
at V except as otherwise noted. In demonstrating compliance, the
D
powerplant or propeller, if retractable, must be retracted,
3.2.19 n —glider negative maneuvering limit load factor
except as otherwise noted.
at V
A
4.3 Load Distribution Limits:
3.2.20 n —glider negative maneuvering limit load factor
4.3.1 Themaximumweightshallbedeterminedsothatitis:
at V
D
2 2 4.3.1.1 Not more than:
3.2.21 q—dynamic pressure = 0.004823 V kg⁄m , when V
(1)The highest weight selected by the applicant, and
is in km/h
(2)The design maximum weight, which is the highest
3.2.22 S—wing area (m )
weight at which compliance with each applicable structural
3.2.23 V—airspeed (m/s, kts)
loadingconditionandallrequirementsforflightcharacteristics
is shown.
3.2.24 V —design maneuvering speed
A
4.3.1.2 Not less than:
3.2.25 V —design cruising speed
C
(1)Forasingle-placeglidernotlessthantheemptyweight
3.2.26 V —design diving speed
D
of the glider, plus a weight of the occupant of 80 kg, plus the
required minimum equipment, plus, for a powered glider,
3.2.27 V —demonstrated flight diving speed
DF
sufficient energy (fuel or other energy storage) for at least 30
3.2.28 V —design flap speed
F
min of flight at maximum continuous power.
3.2.29 V —maximum flap extended speed
FE
(2)Foratwo-placeglidernotlessthantheemptyweightof
3.2.30 V —maximum speed in level flight with maximum the glider, plus a weight of the occupants of 160 kg, plus the
H
continuous power (corrected for sea level standard conditions)
required minimum equipment, plus, for a powered glider,
sufficient energy (fuel or other energy storage) for at least 30
3.2.31 V —maximum speed for landing gear extended
LO
min of flight at maximum continuous power.
3.2.32 V —never exceed speed
NE
4.3.2 The design empty weight shall be specified by the
3.2.33 V —stallingspeedorminimumsteadyflightspeedat
S manufacturer.
which the aircraft is controllable (flaps retracted)
4.3.3 Empty Weight and Center of Gravity Range:
3.2.34 V —stalling speed, or minimum steady flight speed 4.3.3.1 The CG range within which the glider can be safely
S1
in a specific configuration operated must be specified by the manufacturer.
F2564−14 (2022)
TABLE 1 Pilot Force
4.3.3.2 The empty weight, corresponding CG, most
forward,andmostrearwardCGshallbedeterminedwithfixed Wing flaps, landing gear,
air brakes, retraction or
ballast and required minimum equipment.
Pilot force as applied to Pitch, Roll, Yaw,
extension of engine,
the controls N N N
4.3.3.3 The CG range must not be less than that which
two cable release,
N
corresponds to that of a sole pilot weight of 65 kg up to the
For temporary application: 200 150 300 150
maximum weight, always considering the most unfavorable
(less than 2 min) Stick
placing of luggage.
For prolonged application: 20 15 100 Not determined
4.3.3.4 Fixed or removable ballast, or both, may be used if
properly installed and placarded.
4.3.3.5 MultipleESDsmaybeusedifproperlyinstalledand
placarded.
4.5.6.2 Ground roll distance with braking if so equipped.
4.4 Propeller Speed and Pitch Limits for a Powered
4.6 Controllability and Maneuverability:
Glider—Theoperatinglimitationsshallnotallowtheengineto
4.6.1 General:
exceed safe operating limits established by the engine manu-
4.6.1.1 The glider shall be safely controllable and maneu-
facturer under normal conditions.
verable during takeoff, climb, level flight, dive to V or the
DF
4.4.1 Maximum RPM shall not be exceeded with full
maximum allowable speed for the configuration being
throttle during takeoff, climb, or flight at 0.9 V , and 110%
H
investigated,engineextensionandretraction,andapproachand
maximum continuous RPM shall not be exceeded during a
landing through the normal use of primary controls.
glide at V with throttle closed.
NE
4.6.1.2 Smooth transition between all flight conditions shall
4.5 Performance, General—All performance requirements be possible without exceeding pilot force as shown in Table 1.
4.6.1.3 Full control shall be maintained when retracting and
apply in standard ICAO atmosphere in still air conditions and
at sea level. Speeds shall be given in indicated (IAS) and extending flaps within their normal operating speed range (V
S0
to V ).
calibrated (CAS) airspeeds.
FE
4.6.1.4 Lateral,directional,andlongitudinalcontrolshallbe
4.5.1 Stalling Speeds:
possible down to V .
4.5.1.1 Wing level stalling speeds V and V shall be
S0
S0 S
4.6.2 Longitudinal Control:
determined by flight test at a rate of speed decrease of 1 knot/s
4.6.2.1 At steady flight, or if so equipped, with the aircraft
or less, throttle closed, with maximum takeoff weight, and
trimmed as closely as possible for steady flight at 1.3 V ,it
most unfavorable CG.
S1
must be possible at any speed below 1.3 V to pitch the nose
4.5.1.2 For powered gliders, wing level stalling speeds V
S1
S0
downward so that a speed not less than 1.3 V can be reached
and V shall also be determined with the engine idling,
S1
S
promptly. This must be shown with the aircraft in all possible
propeller in the takeoff position, and the cowl flaps closed.
configurations.
4.5.1.3 For powered gliders, wings level, level flight top
4.6.2.2 Longitudinal control forces shall increase with in-
speed V shall be determined by flight test at maximum
H
creasing load factor.
continuous rated RPM or with full throttle, if unable to reach
4.6.2.3 Longitudinal control must be maintained:
max continuous RPM, at maximum takeoff weight, in cruise
(1)In towed flight, while extending or retracting flaps.
configuration.
(2)Whenretractionorextensionoftheairbrakesismadeat
4.5.2 Takeoff for a Powered Glider:
speeds between 1.1 V and 1.5 V .
4.5.2.1 With the glider at maximum takeoff weight and full S0 S0
(3)For powered gliders, when a change of the wing flap
throttle, the distance to clear a 15m (50ft) obstacle shall not
configuration is made during steady horizontal flight at 1.1 V
exceed 600 m (2000 ft). S1
with simultaneous application of maximum continuous power.
4.5.2.2 Takeoff must be demonstrated with crosswind com-
(4)For powered gliders, when the engine is extended or
ponents not less than 0.2 V .
S0
retracted.
NOTE 1—The procedure used for normal takeoff, including flap
4.6.3 Directional and Lateral Control:
position, shall be specified within the AOI.
4.6.3.1 It must be possible, without significant slip or skid,
4.5.3 Climb—At maximum takeoff weight, flaps in the
toreversethedirectionofaturnwitha45°banktotheopposite
position specified for climb within the AOI, landing gear
direction within b/3 or 4 s, whichever is longer (where b is the
retracted, and full throttle, the minimum rate of climb shall
span of the glider in meters), when the turn is made at a speed
exceed 1.0 m/s (200 ft/min).
of 1.4 V , with where applicable, wing flaps, air brakes, and
S1
4.5.4 High Speed Descent—If so equipped, the glider must
landing gear retracted.
not exceed V in a dive at a 30° angle to the horizon with
4.6.3.2 With and without flaps deployed, rapid entry into or
NE
airbrakes extended.
recoveryfromamaximumcross-controlledslipshallnotresult
4.5.5 Descent—If so equipped, the glider must have a glide
in uncontrollable flight characteristics.
slope not flatter than one in seven at a speed of 1.3 V at
4.6.3.3 Lateral and directional control forces shall not re-
S0
maximum weight and with airbrakes extended.
verse with increased deflection.
4.5.6 Landing—The following shall be determined: 4.6.4 Aerotowing:
4.5.6.1 Landing distance from 15 m (50 ft) above ground 4.6.4.1 If the glider is equipped for aerotowing, aerotows
when speed at 15 m (50 ft) is 1.3 V . must be demonstrated at speeds up to V without:
S0 T
F2564−14 (2022)
TABLE 2 Static Longitudinal Stability Requirements
(1)Difficulty in regaining the normal towing position after
the glider has been displaced laterally or vertically. Cruising Configuration
At all speeds between 1.1 V and V
S1 NE
(2)Thereleasedtowcablecontactinganypartoftheglider.
Wing flaps in the position for cruising and for circling
4.6.4.2 Aerotowing must be demonstrated with crosswind
Landing gear retracted
Glider trimmed at 1.4 V and 2 V (if equipped with a trimming device)
components not less than 0.2 V . S1 S1
S0
Air brakes retracted
4.6.4.3 A suitable range of tow cables must be established.
Approach
4.6.4.4 Tests must be repeated for each location of the At all speeds between 1.1 V and V
S1 FE
Wing flaps in the landing position
towing release mechanism.
Landing gear extended
4.6.5 Winch Launching:
Glider trimmed at 1.4 V (if equipped with a trimming device)
S0
Air brakes retracted and extended
4.6.5.1 If the glider is equipped for winch launching or
Climb for Powered Glider
auto-towlaunching,suchlaunchesmustbedemonstratedupto
At all speeds between 0.85 V or 1.05 V
Y Y
V without:
W Wing flaps in the position for climb
Landing gear retracted
(1)Uncontrolled roll after leaving the ground and upon a
Glider trimmed at V (if equipped with a trimming device)
Y
release,
Maximum weight
(2)Uncontrolled pitching oscillations, and
Maximum continuous power
Cruise for Powered Glider
(3)Control forces in excess of those listed in Table 1 and
At all speeds between 1.3 V and V
S1 NE
excessive deflections of the controls.
Wing flaps retracted or in the case of flaps approved for use in cruise
4.6.5.2 Winch launching must be demonstrated with cross- flight in all appropriate positions
Landing gear retracted
wind components not less than 0.2 V .
S0
Glider trimmed for level flight (if equipped with a trimming device)
4.6.5.3 If a trimming device is fitted, the position used
Maximum weight
Power set for horizontal flight at 0.9 V
during the climb must be listed in the AOI. H
Approach for Powered Glider
4.6.6 Approach and Landing:
At all speeds between 1.1 V and V
S1 NE
4.6.6.1 Normal approaches and landings until the glider Wing flaps in the landing position
Landing gear extended
comestoacompletehaltmustbedemonstratedwithcrosswind
Glider trimmed at 1.5 V (if equipped with a trimming device)
S1
components not less than 0.2 V .
S0
Maximum weight
Air brakes retracted and extended
4.6.6.2 Theuseofairbrakesduringapproachwillnotcause
Power set at idle
control forces in excess of those listed in Table 1 or excessive
control displacements, nor affect the controllability of the
glider.
4.6.6.3 After touchdown, there must not be a tendency to
4.6.8 Static Directional and Lateral Stability:
ground loop, for pitching oscillation or to nose over.
4.6.8.1 There can be no tendency for the glider when in
4.6.7 Static Longitudinal Stability:
straight flight at 1.4 V with wing-flaps in all en-route
S1
4.6.7.1 The glider shall demonstrate the ability to trim for
positions, air brakes, and where applicable, landing gear
steady flight at speeds appropriate to the launch, flight, and
retracted to:
landing approach configurations for gliders, and climb and
(1)Turn or bank when the aileron control is released and
cruise for powered gliders; at minimum and maximum weight;
the rudder control held fixed in the neutral position, and
and forward and aft CG limits. If the glider has no in-flight
(2)Yawwhentheruddercontrolisreleasedandtheaileron
adjustable longitudinal trimming device, the trim speed must
control held fixed in the neutral position.
be between 1.2 V and 2.0 V for all CG positions.
S1 S1 4.6.8.2 The glider shall exhibit positive directional and
4.6.7.2 The glider shall exhibit positive longitudinal stabil-
lateralstabilitycharacteristicsatanyspeedabove V ,uptothe
S1
ity characteristics at any speed above V , up to the maximum
maximum allowable speed for the configuration being
S1
allowablespeedfortheconfigurationbeinginvestigated,andat
investigated, and at the most critical CG combination.
the most critical power setting and CG combination.
4.6.8.3 Powered glider must demonstrate:
4.6.7.3 Stability shall be shown by a tendency for the glider
(1)Retractionandextensionofthepowerplantorpropeller
to return toward steady flight after: (1) a “push” from steady
must not produce excessive trim changes,
flightthatresultsinaspeedincrease,followedbyanon-abrupt
(2)A climb at maximum continuous power at V with
Y
release of the pitch control; and (2) a “pull” from steady flight
landing gear retracted and wing flaps in the takeoff position is
that results in a speed decrease, followed by a non-abrupt
achievable with trimmed pitch controls, and
release of the pitch control.
(3)Level flight at all speeds between V and V , with the
Y H
4.6.7.4 The glider shall demonstrate compliance with this landing gear retracted and wing flaps in a position appropriate
section for the conditions listed in Table 2.
to each speed is achievable with trimmed pitch controls.
4.6.7.5 While returning toward steady flight, the aircraft 4.6.8.4 With the glider in straight and steady flight, and
shall: when the aileron and rudder controls are gradually applied in
(1)Not decelerate below stalling speed V , opposite directions, any increase in slideslip angle must corre-
S1
(2)Not exceed V or the maximum allowable speed for spond to an increased deflection of the lateral control. This
NE
the configuration being investigated, and behavior need not follow a linear law.
(3)Exhibitdecreasingamplitudeforanylong-periodoscil- 4.6.9 Dynamic Stability—Anyshortperiodoscillationsshall
lations. be heavily damped within the appropriate speed range (V to
S0
F2564−14 (2022)
V flaps extended and V to V flaps retracted) for primary 4.6.13.2 For gliders in which intentional spinning is
FE S DF
controls fixed and free. In the case of a powered glider, this allowed, the glider must be able to recover from a three-turn
requirement must be met with the engine running at all spin in not more than one and one-half additional turn.
allowable powers. 4.6.13.3 In addition, for either 4.6.13.1 or 4.6.13.2:
(1)The applicable airspeed limit and limit maneuvering
4.6.10 Wings Level Stall:
load factor shall not be exceeded,
4.6.10.1 Itshallbepossibletopreventmorethan30°ofroll
(2)Control forces during the spin or recovery shall not
or yaw by normal use of the controls during the stall and the
exceed those listed in Table 1, and
recovery at all weight and CG combinations.
(3)It must be impossible to obtain unrecoverable spins
4.6.10.2 Thelossofaltitudefromastallmustbedetermined
with any use of the controls.
and listed in the AOI.
4.6.14 Spiral Dive Characteristics—Ifthereisanytendency
4.6.10.3 Minor yaw (up to 5°) shall not have a significant
for a spin to turn into a spiral dive, the glider must be able to
influence on the stall characteristics.
recover from this condition without exceeding either the
4.6.10.4 Compliancewiththissectionmustbedemonstrated
limiting air speed or the limiting maneuvering factor for the
under the following conditions: glider.
(1)Wing flaps in any condition,
4.7 Vibrations—Flight testing shall not reveal by pilot
(2)Air brakes retracted and extended,
observation heavy buffeting (except as associated with a stall),
(3)Landing gear retracted and extended,
excessive airframe or control vibrations, flutter (with proper
(4)Glidertrimmedto1.4 V (ifequippedwithatrimming
S1
attempts to induce it), or control divergence at any speed from
device),
V to V .
S0 DF
(5)Additionally,forpoweredgliders,cowlflapsmustbein
4.8 Ground Control and Stability—There must not be any
the appropriate configuration with the engine at idle and 90%
uncontrollable ground loop tendency at any speed at which a
of maximum continuous power, and
powered glider will operate on the ground up to the maximum
(6)During winch takeoff with the glider pitch 30° above
crosswind component specified in 4.5.2.2.
the horizontal.
4.6.11 Turning Flight Stalls:
5. Structure
4.6.11.1 Whenstalledduringacoordinated45°bankedturn,
5.1 General:
it must be possible to regain normal level flight without
5.1.1 Loads:
encountering uncontrollable rolling or spinning tendencies.
5.1.1.1 Strength requirements are specified in terms of limit
Compliance with this requirement must be shown under the
loads (the maximum loads to be expected in service) and
conditions of 4.6.10.4 that result in the most critical stall
ultimate loads (limit loads multiplied by prescribed factors of
behavior of the glider. The landing configuration, with air-
safety). Unless otherwise provided, prescribed loads are limit
brakes retracted and extended, must be investigated.
loads.
4.6.11.2 Thelossofaltitudefrombeginningofthestalluntil
5.1.1.2 Unless otherwise provided, the air and ground loads
regaining wings level flight and a speed of 1.4 V must be
S1
must be placed in equilibrium with inertia forces, considering
determined.
each item of mass in the aircraft. These loads must be
4.6.12 Stall Warning:
distributed to conservatively approximate or closely represent
4.6.12.1 There must be a clear and distinctive stall warning actual conditions.
with airbrakes, wing flaps, and landing gear in any normal 5.1.1.3 If deflections under load would significantly change
position, both in straight and turning flight. In the case of a the distribution of external or internal loads, this redistribution
must be taken into account.
poweredglider,compliancewiththisrequirementmustalsobe
5.1.2 Factor of Safety:
shown with the engine running in the conditions prescribed in
5.1.2.1 Unless otherwise provided in 5.1.2.2, an ultimate
4.6.10.4(5).
load factor of safety of 1.5 must be used.
4.6.12.2 The stall warning may be furnished either through
5.1.2.2 Special ultimate load factors of safety shall be
the inherent aerodynamic qualities of the glider (that is,
applied according to Table 3.
buffeting) or by a device that will give clearly distinguishable
5.1.3 Strength and Deformation:
indications. A visual only stall warning is not acceptable.
4.6.12.3 The stall warning must begin:
(1)In the speed rand of 1.05 V to 1.1 V ,or
S1 S1
(2)2s to 5 s before the stall occurs while the speed is
TABLE 3 Ultimate Load Factors
decreasing at 1 knot/s.
2.0 × 1.5 = 3.0 on castings
1.2 × 1.5 = 1.8 on fittings
4.6.13 Spinning:
2.0 × 1.5 = 3.0 on bearings at bolted or pinned joints subject to rotation
4.6.13.1 For gliders placarded “no intentional spins,” the
4.45 × 1.5 = 6.67 on control surface hinge-bearing loads except ball and
glider must be able to recover from a one-turn spin or a 3s roller bearing hinges
2.2 × 1.5 = 3.3 on push-pull control system joints
spin, whichever takes longer, in not more than one additional
1.33 × 1.5 = 2 on cable control system joints, seat belt/harness fittings
turn, with the controls used in the manner normally used for
(including the seat if belt/harness is attached to it)
recovery.
F2564−14 (2022)
5.1.3.1 The structure must be able to support limit loads
without permanent deformation.At any load up to limit loads,
the deformation shall not interfere with safe operation.
5.1.3.2 The structure must be able to support ultimate loads
without failure for at least 3 s. However, when proof of
strength is shown by dynamic tests simulating actual load
conditions, the 3s limit does not apply.
5.1.4 Proof of Structure—Each design requirement must be
verified by means of conservative analysis or test (static,
component, or flight), or both.
5.1.4.1 Compliance with the strength and deformation re-
quirements of 5.1.3 must be shown for each critical load
condition. Structural analysis may be used only if the structure
conformstothoseforwhichexperiencehasshownthismethod
to be reliable. In other cases, substantiating load tests must be
made. Dynamic tests, including structural flight tests, are
acceptable if the design load conditions have been simulated.
Substantiating load tests should normally be taken to ultimate
design load.
FIG. 1Maneuvering Envelope
5.1.4.2 Certain parts of the structure must be tested as
specified in 6.11.
5.2 Flight Loads:
5.2.1 General:
5.2.1.1 Flight Load Factors, n, represent the ratio of the
aerodynamic force component (acting normal to the assumed
longitudinal axis of the aircraft) to the weight of the aircraft.A
positive flight load factor is one in which the aerodynamic
force acts upward with respect to the glider.
5.2.1.2 Compliance with the flight load requirements of this
section must be shown at each practicable combination of
weight and disposable load within the operating limitations
specified in the AOI.
5.2.2 Symmetrical Flight Conditions:
5.2.2.1 The appropriate balancing horizontal tail loads must
be accounted for in a rational or conservative manner when
determiningthewingloadsandlinearinertialoadscorrespond-
ing to any of the symmetrical flight conditions specified in
5.2.2 – 5.2.6.
5.2.2.2 The incremental horizontal tail loads due to maneu-
vering and gusts must be reacted by the angular inertia of the
glider in a rational or conservative manner.
FIG. 2Gust Envelope
5.2.2.3 In computing the loads arising in the conditions
prescribed above, the angle of attack is assumed to be changed
5.2.3.1 General—Compliance with the strength require-
suddenly without loss of air speed until the prescribed load
ments of this subpart must be shown at any combination of
factor is attained. Angular accelerations may be disregarded.
airspeed and load factor on and within the boundaries of the
5.2.2.4 The aerodynamic data required for establishing the
flight envelopes specified by the maneuvering and gust criteria
loadingconditionsmustbeverifiedbytests,calculations,orby
of 5.2.3.2 and 5.2.3.3, respectively.
conservative estimation. In the absence of better information,
5.2.3.2 Maneuvering Envelope—Wing flaps are in the en-
the maximum negative lift coefficient for rigid lifting surfaces
route setting and air brakes are closed (see Fig. 1).
shall be assumed to be equal to −0.80. If the pitching moment
5.2.3.3 Gust Envelope—Wing flaps in the en-route setting
coefficient, C , is less than 60.025, a coefficient of at least
mo
(see Fig. 2).At the design maximum speed V , the glider must
D
60.025 must be used.
be capable of withstanding positive (up) and negative (down)
5.2.3 Flight Envelope—Compliance shall be shown at any
gusts of 7.5 m/s acting normal to the flight path.
combination of airspeed and load factor on the boundaries of
5.2.4 Design Airspeeds:
theflightenvelope.Theflightenveloperepresentstheenvelope
5.2.4.1 Design Maneuvering Speed, V :
A
oftheflightloadingconditionsspecifiedbythecriteriaof5.2.4
and 5.2.5 (see Fig. 1). V 5 V =n (1)
A S1 1
F2564−14 (2022)
where: 5.2.7 Unsymmetrical Flight Conditions—The glider is as-
sumedtobesubjectedtotheunsymmetricalflightconditionsof
V = estimated stalling speed at design maximum weight
S1
5.2.7.1 and 5.2.7.2. Unbalanced aerodynamic moments about
with wing-flaps and air brakes retracted, and
the center of gravity must be reacted in a rational or conser-
n = positive limit maneuvering load factor used in design.
vative manner, considering the principle masses furnishing the
5.2.4.2 Design Flap Speed, V —For each landing setting,
F
reacting inertia forces.
V must not be less than the greater of: (1) 1.4 V , where V is
F S S
5.2.7.1 RollingConditions—Theglidershallbedesignedfor
thecomputedstallingspeedwiththewingflapsretractedatthe
the loads resulting from the roll control deflections and speeds
maximumweight;and (2)2.0 V ,where V isthecomputed
SF SF
specified in 5.7.1 in combination with a load factor of at least
stalling speed with wing flaps fully extended at the maximum
twothirdsofthepositivemaneuveringloadfactorprescribedin
weight.
5.2.5.1.
5.2.4.3 Design Aerotow Speed, V ,mustnotbelessthan1.5
T
5.2.7.2 Yawing Conditions—The glider must be designed
V according to 5.2.4.1.
S1
for the yawing loads resulting from the vertical surface loads
5.2.4.4 Design Dive Speed, V :
D
specified in 5.5.
m 1
5.2.8 Loads with Air Brakes and Wing Flaps Extended:
V 518Π~km/h! butnot# V (2)
S DS D
D A
S Cd
min 5.2.8.1 Loads with Air Brakes Extended:
(1)The glider structure must be capable of withstanding
where:
themostunfavorablecombinationofthefollowingparameters:
m
⁄S = wing loading (kg/m ) at design maximum weight,
equivalent air speed at V , air brakes extended, and a load
D
and
factor from 0 to 2.0.
Cd = the lowest possible drag coefficient of the glider.
min
(2)The horizontal tail load corresponds to the static con-
5.2.5 Limit Maneuvering Load Factors:
dition of equilibrium.
5.2.5.1 The positive limit maneuvering load factor n shall
1 (3)In determining the spanwise load distribution, changes
not be less than 4.0 while n shall not be less than 3.0.
inthisdistributionduetothepresenceoftheairbrakesmustbe
5.2.5.2 The negative limit maneuvering load factor n shall
3 accounted for.
not be less than −1.5, while n shall not be less than –2.0.
5.2.8.2 If wing-flaps are installed, positive limit factor 3,0
5.2.6 Gust Load Factors—In the absence of a more rational
must be assumed while positions of the flaps from retracted up
analysis, the gust load factors must be computed as follows:
to positive deflection and up to speed V are considered.
F
k
5.2.8.3 It must be considered that the glider at positions of
ρ UVa
S D
o
the flaps from retracted up to maximum negative deflection
n 516 (3)
mg
must comply with the requirements of 5.2.3.2 and 5.2.3.3.
3 4
S D
S
5.2.9 Engine Torque—The engine mount and its supporting
structure must be designed for the effects of:
where:
3 5.2.9.1 Thelimittorquecorrespondingtotakeoffpowerand
ρ = density of air at sea-level (1225 kg/m ),
o
propeller speed acting simultaneously with 75% of the limit
U = gust velocity (m/s),
loads from flight condition of 5.2.5.1.
V = equivalent air speed (m/s),
5.2.9.2 The limit torque corresponding to maximum con-
a = slope of wing lift curve (1/rad),
tinuous power and propeller speed acting simultaneously with
m = mass of the glider (kg),
g = acceleration due to gravity (m/s ), the limit loads from the flight condition of 5.2.5.1.
S = wing area (m ), and
5.2.9.3 Forconventionalreciprocatingengineswithpositive
k = gust alleviation factor calculated from the following
drive to the propeller, the limit torque to be accounted for in
formula:
5.2.9.1 and 5.2.9.2 is obtained by multiplying the mean torque
by one of the following factors:
0.88µ
k 5 (4)
(1)2 for engines with 4 cylinders,
5.31µ
(2)3 for engines with 3 cylinders,
where:
(3)4 for engines with 2 cylinders, and
(4)8 for an engine with one cylinder.
m
5.2.9.4 For conventional electric motors with positive drive
S
µ 5 non 2dimensionalglidermassration (5)
~ !
ρCa to the propeller, the limit torque to be accounted for in 5.2.9.1
and5.2.9.2isobtainedbymultiplyingthemeantorqueby1.33.
where:
5.2.10 Side Load on Engine Mount:
ρ = density of air (kg/m ) at the sea level, and
5.2.10.1 Theenginemountanditssupportingstructuremust
C = mean geometric chord of wing (m).
be designed for a limit load factor in a lateral direction, for the
The value of n calculated from the expression given above
side load on the engine mount, of not less than one third of the
need not exceed: 1
limit load factor for flight condition A of Fig. 1 ( ⁄3 n ).
V 5.2.10.2 The side load prescribed in 5.2.10.1 shall be
n 5 (6)
S D
V assumed to be independent of other flight conditions.
S1
F2564−14 (2022)
5.3 Control Surface and System Loads: 5.3.8.2 Aforceequalto12timestheweightappliedforeand
aft and parallel to the hinge line.
5.3.1 Control Surface Loads—The control surface loads
5.3.9 The motion of wing flaps on opposite sides of the
specified in 5.3.3 through 5.3.7 are assumed to occur in the
conditions described in 5.2.2 through 5.2.6. plane of symmetry must be synchronized by a mechanical
interconnectionunlesstheaircrafthassafeflightcharacteristics
5.3.2 Control System Loads—Each part of the primary
with the wing flaps retracted on one side and extended on the
control system situated between the stops and the control
other.
surfaces must be designed for the loads corresponding to at
5.3.10 All primary controls shall have stops within the
least 125% of the of the computed hinge moments of the
systemtowithstandthegreaterofpilotforce,125%ofsurface
movable control surfaces resulting from the loads in the
loads, or ground gust loads (see 5.3.7).
conditions prescribed in 5.3.1 through 5.7.3. In computing the
hinge moments, reliable aerodynamic data must be used. In no
5.4 Horizontal Stabilizing and Balancing Surfaces:
case shall the load in any part of the system be less than those
5.4.1 Balancing Loads:
resulting from the application of 60% of the pilot forces
5.4.1.1 Ahorizontalstabilizingsurfacebalancingloadifthe
described in 5.3.3. In addition, the system limit loads need not
load necessary to maintain equilibrium in any specified flight
exceed the loads that can be produced by the pilot. Pilot forces
condition with no pitching acceleration.
used for design need not exceed the maximum pilot forces
5.4.1.2 Horizontal stabilizing surfaces must be designed for
prescribed in 5.3.3.
the balancing loads occurring at any point on the limit
5.3.3 Loads Resulting from Limit Pilot Forces:
maneuvering envelope and in the air-brake and wing-flap
5.3.3.1 The main control systems for the direct control of
positions specified in 5.2.3.
the aircraft about its longitudinal, lateral, or yaw axis, includ-
5.4.2 Maneuvering Loads—Horizontal stabilizing surfaces
ing the supporting points and stops, must be designed for the
must be designed for pilot-induced pitching maneuvers im-
limit loads resulting from the limit pilot forces given in Table
posed by the following conditions:
1.
5.4.2.1 At speed V , maximum upward deflection of pitch
A
5.3.3.2 The rudder control system must be designed to a
control surface,
load of 600 N per pedal acting simultaneously on both pedals
5.4.2.2 At speed V , maximum downward deflection of
A
in the forward direction.
pitch control surface,
5.3.4 Dual-Control Systems—Dual-control systems must be
5.4.2.3 At speed V , one-third maximum upward deflection
D
designed for the loads resulting from each pilot applying 0.75 of pitch control surface, and
times the load specified in 5.3.3 with the pilots acting in
5.4.2.4 At speed V , one-third maximum downward deflec-
D
opposition. tion of pitch control surface.
5.3.5 Secondary Control Systems—Secondary control
NOTE2—In5.4.2,thefollowingassumptionsshouldbemade:theglider
systems, such as those for landing gear retraction or extension,
is initially in level flight, and its altitude and airspeed do not change. The
wheelbrake,trimcontrol,andsoforthmustbedesignedforthe
loads are balanced by inertia forces.
maximum forces that a pilot is likely to apply.
5.4.3 Gust Loads—In the absence of a more rational
5.3.6 Control System Stiffness and Stretch—The amount of
analysis,thehorizontaltailloadsmustbecomputedasfollows:
controlsurfaceortabmovementavailabletothepilotshallnot
ρ dε
o
be dangerously reduced by elastic stretch or shortening of the
F 5 F 1 S a UkH 1 2 V (8)
S D
VOP o VOP VOP VOP
2 dα
system in any condition.
5.3.7 Ground Gust Conditions—In the absence of a more
where:
rational analysis, the control system from the control surfaces
F = horizontal tail balancing load acting on the hori-
o
to the stops or control locks, when installed, must be designed
zontal tail before the appearance of the gust (N),
for limit loads due to gusts corresponding to the following
ρ = density of air at sea-level (1225 kg/m ),
o
hinge moments:
S = area of horizontal tail (m ),
VOP
a = slope of horizontal tail lift curve per radian,
VOP
M 5 k·C ·S ·q (7)
S S S
U = gust speed (m/s),
where:
kH = gustfactor.Intheabsenceofarationalanalysis,the
VOP
same value shall be taken as for the wing,
M = limit hinge moment,
S
V = speed of flight (m/s), and
C = mean chord of the control surface aft of the hinge line,
S
dε
S = area of the control surface aft of the hinge line, ⁄dα = rateofchangeofdownwashanglewithwingangle
S
Q = dynamic pressure corresponding to an airspeed of 38 of attack.
knots, and
5.5 Vertical Stabilizing Surfaces:
K = limit hinge moment coefficient due to ground gust =
5.5.1 Maneuvering Loads—The vertical stabilizing surfaces
0.75.
must be designed for maneuvering loads imposed by the
5.3.8 Control Surface Mass Balance Weights—Ifapplicable,
following conditions:
shall be designed for the following forces to be applied to the 5.5.1.1 At a speed, the greater of V and V , full deflection
A T
mass balance weight:
of the rudder.
5.5.1.2 At speed V , one-third full deflection of the rudder.
5.3.8.1 A force equal to 24 times the mass balance weight
D
applied normal to the surface, and 5.5.2 Gust Loads:
F2564−14 (2022)
5.5.2.1 Theverticalstabilizingsurfacesmustbedesignedto of the surfaces at speed higher than V . This condition is
A
withstand lateral gusts of the values prescribed in 5.2.3.3. supplemental to the equivalent horizontal and vertical cases
previously specified.
5.5.2.2 In the absence of a more rational analysis, the
vertical surfaces gust loads shall be computed as follows:
5.7 Ailerons, Wing Flaps, and Special Devices:
ρ
5.7.1 Ailerons—The ailerons must be designed for control
o
F 5 a S UkV (9)
SOP SOP SOP
loads corresponding to the following conditions:
5.7.1.1 At speed V , the full deflection of the roll control.
where: A
5.7.1.2 At speed V , one-third of the full deflection of the
T
F = gust load (N),
SOP
roll control.
a = slope of vertical tail lift curve per radian,
v
S = area of vertical tail (m ), 5.7.2 Flaps—Wing flaps, their operating mechanisms, and
SOP
ρ = density of air at sea-level (1.225 kg/m ),
supporting structure must be designed for the critical loads
o
V = speed of flight (m/s),
occurring in the flaps-extended operating range with the flaps
U = gust speed (m/s), and
in any position.
k = gust factor, could be taken as 1.2.
5.7.3 Special Devices—The loadings for special devices
using aerodynamic surfaces, such as air brakes, must be
5.5.3 Outboard Fins or Winglets:
determined from test data or reliable aerodynamic data that
5.5.3.1 If outboard fins or winglets are on the horizontal
allows close estimates.
surfaces or wings, the horizontal surfaces or wings must be
designed for their maximum load in combination with loads
5.8 Ground Load Conditions:
induced by the fins or winglets and moments or forces exerted
5.8.1 Basic Landing Conditions—The limits of the ground
on the horizontal surfaces or wings by the fins or winglets.
loads specified in this subpart are considered to be external
5.5.3.2 If outboard fins or winglets extend above and below
loadsandinertialforcesthatactuponagliderstructure.Ineach
the horizontal surface, the critical vertical surface loading (the
specified ground load condition, the external reactions must be
load per unit area determined in accordance with 5.5.1 and
placedinequilibriumwiththelinearandangularinertialforces
5.5.2) must be applied to:
in a rational or conservative manner. At the design maximum
(1)The part of the vertical surface above the horizontal
weight, the selected limit of the vertical inertia load factor at
surface with 80% of that loading applied to the part below the
theCGofthegliderforthegroundloadconditionsshallnotbe
horizontal surface or wing, and
less than that which would be obtained when landing with a
(2)The part of the vertical surface below the horizontal
descent velocity of 1.5 m/s. Wing lift balancing the weight of
surface or wing with 80% of that loading applied to the part
the glider shall be assumed to act through the CG. The ground
above the horizontal surface or wing.
reaction load factor shall be equal to the inertia load factor
5.5.3.3 The end plate effects of outboard fins or winglets minus one.
must be taken into account in applying the yawing conditions
5.8.2 Subsections 5.8.3 through 5.8.8 apply to a glider with
of 5.5.1 and 5.5.2 to the vertical surfaces in 5.5.3.2.
conventionalarrangementsoflandinggear.Forunconventional
5.5.3.4 When rational methods are used for computing types, it may be necessary to investigate additional landing
loads, the maneuvering loads of 5.5.1 on the vertical surfaces
conditions, depending on the arrangement and design of the
and the n = 1 horizontal surface or wing load, including landing gear units.
induced loads on the horizontal surface, or wing and moments
5.8.3 Level Landing Conditions:
or forces exerted on the horizontal surfaces or wing, must be
5.8.3.1 Foralevellanding,thegliderisassumedtobeinthe
applied simultaneously for the structural loading condition.
following attitudes:
(1)For gliders with a tail skid or wheel, or both, a normal
5.6 Supplementary Conditions for Stabilizing Surfaces:
level flight attitude.
5.6.1 Combined Loads on Stabilizing Surfaces:
(2)For gliders with nose wheels, attitudes in which the
5.6.1.1 With the aircraft in a loading condition correspond-
nose and main wheels contact the ground simultaneously; and
ingtoAorDinFig.1(whicheverconditionleadstothehigher
the main wheels contact the ground and the nose wheel is just
balance load) the loads on the horizontal surface must be
clear of the ground.
combined with those on the vertical surface as specified in
5.8.3.2 The main gear vertical load component F must be
V
5.5.1. It must be assumed that 75% of the loads according to
determined to the conditions in 6.12.3.
5.4.2 for the horizontal stabilizing surface and 5.5.1 for the
5.8.3.3 The main gear vertical load component F must be
vertical stabilizing surface are acting simultaneously. V
combined with a rearward acting horizontal component F so
H
5.6.1.2 The stabilizing surfaces and fuselage must be de-
that the resultant load acts at an angle at 30° with the vertical.
signed for asymmetric loa
...