ASTM F2317/F2317M-16a

This document is not an ASTM standard and is intended only to provide the user of an ASTM standard an indication of what changes have been made to the previous version. Because
it may not be technically possible to adequately depict all changes accurately, ASTM recommends that users consult prior editions as appropriate. In all cases only the current version
of the standard as published by ASTM is to be considered the official document.
Designation: F2317/F2317M − 16 F2317/F2317M − 16a
Standard Specification for
Design of Weight-Shift-Control Aircraft
This standard is issued under the fixed designation F2317/F2317M; the number immediately following the designation indicates the year
of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval.
A superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope
1.1 This specification covers the minimum airworthiness standards a manufacturer shall meet in the designing, testing, and
labeling of weight-shift-control aircraft.
1.2 This specification covers only weight-shift-control aircraft in which flight control systems do not use hinged surfaces
controlled by the pilot.
NOTE 1—This section is intended to preclude hinged surfaces such as typically found on conventional airplanes such as rudders and elevators. Flexible
sail surfaces typically found on weight-shift aircraft are not considered hinged surfaces for the purposes of this specification.
1.3 Weight-shift-control aircraft means a powered aircraft with a framed pivoting wing and a fuselage (trike carriage)
controllable only in pitch and roll by the pilot’s ability to change the aircraft’s center of gravity with respect to the wing. Flight
control of the aircraft depends on the wing’s ability to flexibly deform rather than the use of control surfaces.
1.4 This specification is organized and numbered in accordance with the bylaws established for Committee F37. The main
sections are:
Scope 1
Referenced Documents 2
Terminology 3
Flight Requirements 4
Structural Requirements 5
Design and Construction Requirements 6
Powerplant Requirements 7
Equipment Requirements 8
Operating Limitations 9
Keywords 10
Annex Annex A1
Appendix Appendix X1
1.5 The values stated in either SI units or inch-pound units are to be regarded separately as standard. The values stated in each
system may not be exact equivalents; therefore, each system shall be used independently of the other. Combining values from the
two systems may result in non-conformance with the standard.
1.6 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 and health practices and determine the applicability of regulatory
requirements prior to use.
2. Referenced Documents
2.1 ASTM Standards:
F2339 Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft
F2506 Specification for Design and Testing of Light Sport Aircraft Propellers
2.2 Federal Aviation Regulations:
FAR-33 Airworthiness Standards: Aircraft Engines
FAR-35 Airworthiness Standards: Propellers
This specification is under the jurisdiction of ASTM Committee F37 on Light Sport Aircraft and is the direct responsibility of Subcommittee F37.40 on Weight Shift.
Current edition approved June 1, 2016Nov. 1, 2016. Published July 2016December 2016. Originally approved in 2005. Last previous edition approved in 20102016 as
F2317/F2317M – 10.F2317/F2317M – 16. DOI: 10.1520/F2317_F2317M-16.10.1520/F2317_F2317M-16A.
For referenced ASTM standards, visit the ASTM website, www.astm.org, or contact ASTM Customer Service at service@astm.org. For Annual Book of ASTM Standards
volume information, refer to the standard’s Document Summary page on the ASTM website.
Available from Federal Aviation Administration, 800 Independence Ave., SW, Washington, DC 20591.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F2317/F2317M − 16a
2.3 Joint Aviation Requirements:
JAR-E Engines
JAR-P Propellers
JAR-22 Sailplanes and Powered Sailplanes
3. Terminology
3.1 Definitions—Aircraft Weight:
3.1.1 design maximum aircraft weight, n—aircraft design maximum weight W shall be the sum of W + W .
MAX WING SUSP
3.1.2 design maximum trike carriage weight, n—design maximum trike carriage weight, W , shall be established so that it is:
susp
(1) highest trike carriage weight at which compliance with each applicable structural loading condition and each applicable flight
requirement is shown, and (2) not less than the empty trike carriage weight, W , plus a weight of occupant(s) of 86.0 kg [189.6
tkmt
lb] for a single-seat aircraft or 150 kg [330.8 lb] for a two-seat aircraft, plus the lesser of full usable fuel or fuel weight equal to
1-h burn at economical cruise at maximum gross weight.
3.1.3 trike carriage empty weight, W , n—all parts, components, and assemblies that comprise the trike carriage assembly or
tkmt
that are attached to the suspended trike in flight, including any wing attachment bolts, shall be included in the trike carriage
assembly empty weight, W . These must include the required minimum equipment, unusable fuel, maximum oil, and where
tkmt
appropriate, engine coolant and hydraulic fluid. Trike carriage empty weight, W , shall be recorded in the Aircraft Operating
tkmt
Instructions (AOI).
3.1.4 wing weight, W , n—all parts, components, and assemblies that comprise the wing assembly, or that are attached to the
wing
wing in flight, shall be included in the wing weight, W . The wing weight, W , shall be entered in the AOI.
wing wing
3.2 Abbreviations:
3.2.1 AOI—Aircraft Operating Instructions
3.2.2 C—Celsius
3.2.3 CAS—calibrated air speed (m/s, kts)
3.2.4 cm—centimetre
3.2.5 daN—deca Newton
3.2.6 F—Fahrenheit
3.2.7 Hg—mercury
3.2.8 IAS—indicated air speed (m/s, kts)
3.2.9 in.—inch
3.2.10 ISA—international standard atmosphere
3.2.11 kg—kilogram
3.2.12 kt(s)—nautical mile per hour (knot) (1 nautical mph = (1852 ⁄3600) m/s)
3.2.13 lb—pound (1 lb = 0.4539 kg)
3.2.14 m—metre
3.2.15 mb—millibars
3.2.16 N—Newton
3.2.17 psi—pounds per square inch gage pressure
3.2.18 s—seconds
3.2.19 SI—international system of units
3.2.20 V —design maneuvering speed
A
3.2.21 V —design cruising speed
C
3.2.22 V —demonstrated flight diving speed
DF
3.2.23 V —maximum sustainable speed in straight and level flight
H
3.2.24 V —never exceed speed
NE
3.2.25 V —stalling speed or minimum steady flight speed at which the aircraft is controllable in the landing configuration
S0
3.2.26 V —stalling speed, or the minimum steady flight speed in a specific configuration
S1
3.2.27 V —speed for best angle of climb
x
Available from Global Engineering Documents, 15 Inverness Way, East Englewood, CO 80112-5704
F2317/F2317M − 16a
3.2.28 V —speed for best rate of climb
y
3.2.29 V —maximum aerotow speed
T
3.2.30 W —maximum design weight
MAX
3.2.31 WSC—weight shift control (aircraft)
4. Flight Requirements
4.1 Proof of Compliance:
4.1.1 It shall be possible to demonstrate that the aircraft meets the requirements in this section at each allowable combination
of weight, hang point, and trimmer setting.
4.1.2 The test aircraft used to demonstrate compliance with this specification shall be an accurate representation of the
production aircraft except in the following case:
4.1.2.1 For the purposes of this test only, the aircraft may be modified to expand the control travel or limits in pitch when
establishing V or V .
DF S1
4.1.3 Airspeeds shall be corrected to standard atmospheric conditions 1013.25 mb [29.92 in. Hg], 15°C [59°F].
4.1.4 Climb performance requirements shall be met at standard conditions or conditions more adverse.
4.2 General Performance:
4.2.1 Stall Speed in the Landing Configuration (V ):
S0
4.2.1.1 The stall speed, if obtainable, or the minimum flight speed shall be established with: (1) engine idling with the throttle
closed, (2) hang point that produces the highest stalling or minimum flight speed, (3) maximum takeoff weight, and (4) trim setting
in the landing configuration.
4.2.1.2 V shall be determined by flight-testing, in accordance with the following procedures: (1) aircraft power at idle, at a
S0
speed of not less than V plus 2.6 m/s [5 kts], and (2) the speed reduced at a rate not exceeding 0.5 m/s [1 kt/s] until the stall is
S0
produced as indicated by an autonomous downward pitching motion of the wing or until the control limit is reached.
4.2.1.3 It shall be possible to prevent more than 30° of roll or yaw by normal use of the controls during the stall and the recovery,
or, if stall is not achieved before the control limit is reached, during the slowing to V and subsequent acceleration to V plus
S0 S0
2.6 m/s [5 kts].
4.2.2 Stall Speed Free of Control Limits (V ):
S1
4.2.2.1 Where control limits result in V being reached before the aircraft stalling, then the stall speed free of control limits
S0
(V ) shall be determined. V shall be established with: (1) the aircraft in the landing configuration defined in 4.2.1.1, and (2) the
S1 S1
aircraft may be modified for the purposes of this test, only to expand the nose up pitch control range to the extent necessary for
the aircraft to stall when flown in accordance with the procedures detailed in 4.2.1.2.
4.2.2.2 Where V as determined in accordance with the procedures of 4.2.1.2 is the speed at which the aircraft stalls, then V
S0 S1
= V .
S0
4.2.3 Minimum Climb Performance:
4.2.3.1 The gradient of climb at recommended takeoff power at Vx shall not be less than 1:12.
4.2.3.2 The rate of climb shall exceed 1.5 m/s [300 ft/min] at Vy at recommended takeoff power.
4.2.4 Flutter, Buffeting, and Vibration—Flight-testing shall not reveal, by pilot observations, potentially damaging buffeting,
airframe, or controls vibration, flutter (with attempts to induce it), or control divergence, at any speed from V to V .
S0 DF
4.2.5 Turning Flight and Stalls—Stalls shall be performed as follows: after establishing a steady state turn of at least 30° bank,
the speed shall be reduced until the aircraft stalls, or until the full nose up limit of pitch control is reached. After the turning stall
or reaching the limit of pitch control, level flight shall be regained without exceeding 60° of roll. This shall be performed with the
engine at idle. No loss of altitude greater than 152 m [500 ft], uncontrolled turn of more than one revolution, or speed buildup to
greater than V shall be associated with the recovery.
NE
4.2.6 V —Maximum sustainable speed in straight and level flight, knots CAS.
H
4.2.6.1 V shall be established in straight and level flight with: (1) maximum allowed continuous engine power, and (2) the
H
combination of weight, loading, trimmer setting, and use of the flight controls allowed by the manufacturer that yields the highest
sustainable speed.
NOTE 2—In the case where maximum continuous engine power results in a climb at maximum speed, power may be reduced as needed to maintain
level flight.
4.3 Controllability and Maneuverability:
4.3.1 General—When operating in accordance with the recommendations in the Aircraft Operating Instructions, the aircraft
shall be safely controllable and maneuverable during:
4.3.1.1 Takeoff at maximum takeoff power,
4.3.1.2 Climb,
4.3.1.3 Level flight,
4.3.1.4 Descent,
4.3.1.5 Landing, power on and off,
F2317/F2317M − 16a
4.3.1.6 With sudden engine failure,
4.3.1.7 Turns,
4.3.1.8 Changing speeds between V and V , and
S0 NE
4.3.1.9 Dive to V .
NE
4.3.2 Longitudinal Control:
4.3.2.1 Starting at a speed of 1.1 V , it shall be possible to pitch the nose downwards so that a speed equal to 1.3 V can be
S0 S0
reached in less than 4 s.
4.3.2.2 It shall be possible to pitch the nose up at V at the most adverse hang point, trimmer setting, and engine power.
NE
4.3.3 Lateral Control:
4.3.3.1 Using an appropriate control action, it shall be possible to reverse a steady 30° banked turn to a 30° banked turn in the
opposite direction. This shall be possible in both directions within 5 s from initiation of roll reversal, with the aircraft flown at 1.3
V .
S0
4.3.3.2 Lateral control forces shall not reverse with increased displacement of the flight controls.
4.3.4 Trim Speeds—The speeds to achieve longitudinal trim shall lie between 1.3 V and 0.909 V at all engine powers and
S0 NE
the allowable hang points.
4.3.5 Ground Handling—It shall be possible to prevent ground looping, with normal use of controls, up the maximum crosswind
component published in the AOI.
4.4 Stability:
4.4.1 Longitudinal Stability:
4.4.1.1 The aircraft shall demonstrate the ability to sustain steady flight at speeds appropriate for climb, cruise, and landing.
4.4.1.2 A pull force shall be required to attain and maintain any speed above trim and a push force shall be required to attain
and maintain any speed below trim. As the control force is reduced, the aircraft shall return to within 20 % the original trim speed.
4.4.2 Pitch Testing—A test of the wing pitching moment about the hang point shall be conducted at V × 0.866 over the range
S0
of angles of attack from 15° above zero lift angle to 10° below zero lift angle of attack. The wing shall exhibit a trim angle above
zero lift angle of attack, and a positive pitching moment at any angle below trim, or if trim is not achieved in the test range, the
wing shall exhibit a positive pitching moment throughout the range of angles specified.
NOTE 3—This test may be conducted as a taxi test with the wing mounted to the trike carriage.
5. Structural Requirements
5.1 Strength Requirements:
5.1.1 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 as specified in 5.3). Unless otherwise provided, prescribed loads are
limit loads.
5.1.2 The structure shall be able to support limit loads without permanent deformation. At any load up to limit loads, the
deformation may not interfere with safe operation.
5.1.2.1 The structure shall be able to support ultimate loads with a positive margin of safety (analysis) or without failure for
at least 3 s (tests).
5.2 Fulfillment of Design Requirements:
5.2.1 Fulfillment of the design requirements shall be determined by conservative analysis, tests, or a combination of both.
Structural analysis alone may be used for validation of the structural requirements only if the structure conforms to those for which
experience has shown this method to be reliable. Aerodynamic data required for the establishment of the loading conditions shall
be verified by tests, calculations, or conservative estimation.
5.2.1.1 For analysis and test purposes, unless otherwise provided, the air and ground loads shall be placed in equilibrium with
inertia forces, considering each major item of mass in the aircraft. The loads shall be distributed so as to represent actual conditions
or a conservative approximation to them.
5.2.2 If deflections under load would significantly change the distribution or amount of external or internal loads, this
redistribution shall be taken into account.
5.2.3 The results obtained from strength tests should be corrected for departures from the minimal mechanical material
properties and least favorable material dimensional tolerance values defined in the design.
5.3 Safety Factors—The factor of safety is 1.5, except it shall be increased to:
F2317/F2317M − 16a
3 on castings and bearings whose failure would
preclude continued safe flight and landing of
the aircraft or result in serious injury to the
occupants
2 on other castings and bearings
2 on cables
2 on lap belts and shoulder harnesses
1.73 on fittings and system joints whose strength is
not proven by limit and ultimate tests in which
actual stress conditions apply or are simulated.
5.4 Design Airspeeds:
5.4.1 The selected design air speeds are calibrated air speeds (CAS):
5.4.1.1 Maneuvering Speed V —V shall be greater than or equal to V × 2.
A A S1
5.4.2 V shall be no greater than 0.9 × V .
NE DF
5.4.3 V shall be greater than or equal to the lesser of 1.11 × V or 1.11 × V .
DF A DMAX
5.5 Flight Loads:
5.5.1 Except in the case of dynamic testing, as detailed in the applicable sections of this specification, the limit load factors must
have at least the following values:
+4.0
−2.0
5.5.1.1 If V is greater than two times V , then the minimum positive limit load factor shall equal (V /V ) . The negative load
A S1 A S1
limit factor shall not be required to be greater than −2.0.
5.5.2 Although it is very difficult and very unlikely to achieve sustained negative flight loads on weight shift aircraft, wings shall
be tested for such loads to ensure adequate strength to withstand negative loads caused by gusts, landing, and ground handling.
5.5.3 Adequate structure of the wing to ultimate loads as prescribed by the 1.5 safety factor shall be verified via test (static,
component, dynamic, or flight).
5.5.3.1 Compliance with special factors above the safety factor of 1.5 may be shown by testing or conservative analysis, or both.
5.5.3.2 For a conventional flex-wing configuration, for the purposes of calculating the positive and negative limit and ultimate
load values for test purposes, unless a specific testing protocol listed in this specification or its appendices is used that specifies
another method for allocating the weight of the wing, it shall be considered appropriate to include in the weight of the aircraft 50 %
of the weight of all components comprising the wing assembly.
5.5.4 For static testing of the wing, in the absence of a more rational analysis, the test shall be conducted in accordance with
one of the test protocols as contained in Appendix X1.
5.5.5 Compliance with the positive limit load requirements for the wing may alternatively be shown by a dynamic test of the
wing in which the wing is tested at an angle of attack equal to the highest angle at which maximum lift is achieved, at an airspeed
equal to the greater of 1.0 × V (maneuvering speed), or the speed that will produce a measured load of 3.8 Gs, for a minimum
A
of 3 s without permanent deformation of the structure.
5.5.6 Compliance with the positive ultimate load requirements for the wing may alternatively be shown by a dynamic test of
the wing in which the wing is tested at an angle of attack equal to the highest angle at which maximum lift is achieved, at an
airspeed equal to the greater of 1.225 × V (maneuvering speed), or the speed which produces 1.5 times the load achieved in the
A
limit load test, for a minimum of 3 s without failure.
5.5.7 Compliance with the negative limit load requirements for the wing may alternatively be shown by a dynamic test of the
wing in which the wing is tested at a negative angle of attack equal to the highest negative angle at which maximum negative lift
is achieved, at an airspeed equal to the greater of 0.707 × V , or the speed which produces a measured negative load of 1.52 Gs,
A
for a minimum of 3 s without permanent deformation of the structure.
5.5.8 Compliance with the negative ultimate load requirements for the wing may alternatively be shown by a dynamic test of
the wing in which the wing is tested at a negative angle of attack equal to the highest negative angle at which maximum negative
TABLE 1 Pilot Forces
Pilot Force, Method of
Control
daN [lb force] Force Application
Pitch 66.7 [150] Push or pull of control bar
Roll 31.1 [70] Lateral force (roll) applied
to the control bar
Foot controls for 89 [200] Apply forward pressure
steering on one pedal
Foot controls for 44.5 [100] Push of control
throttle and brake
Miscellaneous 22.2 [50] Push and pull of control
secondary controls lever
F2317/F2317M − 16a
lift is achieved, equal to the greater of 0.866 × V (maneuvering speed), or the speed which produces 1.5 times the load achieved
A
in the limit load test, for a minimum of 3 s without failure.
5.5.9 If dynamic testing is chosen for limit load testing of the wing, compliance with the ultimate load requirements may be
shown by conducting a static load test to a load of 1.5 times the loads generated during dynamic limit tests. The wing shall sustain
this load for a minimum of 3 s without failure but may show permanent deformation.
5.6 Pilot Control Loads:
5.6.1 The pitch and roll control bar shall be designed to withstand as far as to the stops (these included) limit loads arising from
the pilot forces in Table 1. Lower pilot forces may be established, provided it can be demonstrated that the forces in Table 1 are
unlikely to be able to be applied.
5.6.1.1 Where a backup safety system ensures the ability to continue safe flight in the event of a control system component
failure, the forces in Table 1 may be reduced by ⁄3 .
5.6.1.2 In roll, in the case in which the rear lower rigging wires bearing against the operators or trike fuselage is the only
practicable roll control limit stop preventing damage to the structure, the control frame upright shall be able to achieve an angle
within 10º of the vertical centerline of the trike without causing permanent structural deformation. If the upright can reach this
angle, it is not necessary to show compliance with Table 1 for control stop purposes.
5.6.2 Dual control systems must be designed to withstand the loads that result when each pilot applies 0.75 times the load
specified in Table 1 with the pilots acting together in the same direction, and the pilots acting in opposition.
5.7 Ground Loads:
5.7.1 The fuselage shall be able to sustain a static limit load of 2g without permanent deformation. The loads shall be distributed
throughout the structure in a rational manner, including wing load, engine load, full fuel load, occupant load, frame load, and
maximum allowable baggage load.
5.7.2 An ultimate load of 2 g × 1.5 safety factor (3 g) shall be supported for a minimum of 3 s without failure.
5.7.3 Landing Gear Shock Absorption—The landing gear shall be capable of absorbing the energy that would resu
...