ASTM F2317 F2317M-2016 Standard Specification for Design of Weight-Shift-Control Aircraft《航空器重量变化控制设计的标准规格》.pdf

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1、Designation: F2317/F2317M 10F2317/F2317M 16Standard Specification forDesign of Weight-Shift-Control Aircraft1This standard is issued under the fixed designation F2317/F2317M; the number immediately following the designation indicates the yearof original adoption or, in the case of revision, the year

2、 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. Scope1.1 This specification covers the minimum airworthiness standards a manufacturer shall meet in the designing, testing,

3、 andlabeling of weight-shift-control aircraft.1.2 This specification covers only weight-shift-control aircraft in which flight control systems do not use hinged surfacescontrolled by the pilot.NOTE 1This section is intended to preclude hinged surfaces such as typically found on conventional airplane

4、s such as rudders and elevators. Flexiblesail 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

5、 only in pitch and roll by the pilots ability to change the aircrafts center of gravity with respect to the wing. Flightcontrol of the aircraft depends on the wings ability to flexibly deform rather than the use of control surfaces.1.4 This specification is organized and numbered in accordance with

6、the bylaws established for Committee F37. The mainsections are:Scope 1Referenced Documents 2Terminology 3Flight Requirements 4Structural Requirements 5Design and Construction Requirements 6Powerplant Requirements 7Equipment Requirements 8Operating Limitations 9Keywords 10Annex Annex A1Appendix Appen

7、dix X11.5 The values stated in either SI units or inch-pound units are to be regarded separately as standard. The values stated in eachsystem may not be exact equivalents; therefore, each system shall be used independently of the other. Combining values from thetwo systems may result in non-conforma

8、nce 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 responsibilityof the user of this standard to establish appropriate safety and health practices and determine the applicability of regulatoryrequirements prior t

9、o use.2. Referenced Documents2.1 ASTM Standards:2F2339 Practice for Design and Manufacture of Reciprocating Spark Ignition Engines for Light Sport Aircraft2.2 Federal Aviation Regulations:3FAR-33 Airworthiness Standards: Aircraft EnginesFAR-35 Airworthiness Standards: Propellers2.3 Joint Aviation Re

10、quirements:4JAR-E EnginesJAR-P Propellers1 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 Jan. 1, 2010June 1, 2016. Published February 2010July 2016. Originall

11、y approved in 2005. Last previous edition approved in 20092010 asF2317/F2317M 05 (2009).F2317/F2317M 10. DOI: 10.1520/F2317_F2317M-10.10.1520/F2317_F2317M-16.2 For referencedASTM standards, visit theASTM website, www.astm.org, or contactASTM Customer Service at serviceastm.org. For Annual Book of AS

12、TM Standardsvolume information, refer to the standards Document Summary page on the ASTM website.3 Available from Federal Aviation Administration, 800 Independence Ave., SW, Washington, DC 20591.4 Available from Global Engineering Documents, 15 Inverness Way, East Englewood, CO 80112-5704This docume

13、nt 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. Becauseit may not be technically possible to adequately depict all changes accurately, ASTM recommends that users consult prior editions as ap

14、propriate. In all cases only the current versionof the standard as published by ASTM is to be considered the official document.Copyright ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States1JAR-22 Sailplanes and Powered Sailplanes3. Terminology3.1 D

15、efinitionsAircraft Weight:3.1.1 design maximum aircraft weight, naircraft design maximum weight WMAX shall be the sum of WWING + WSUSP.3.1.2 design maximum trike carriage weight, ndesign maximum trike carriage weight, Wsusp, shall be established so that it is:(1) highest trike carriage weight at whi

16、ch compliance with each applicable structural loading condition and each applicable flightrequirement is shown, and (2) not less than the empty trike carriage weight, Wtkmt, plus a weight of occupant(s) of 86.0 kg 189.6lb for a single-seat aircraft or 150 kg 330.8 lb for a two-seat aircraft, plus th

17、e lesser of full usable fuel or fuel weight equal to1-h burn at economical cruise at maximum gross weight.3.1.3 trike carriage empty weight, Wtkmt, nall parts, components, and assemblies that comprise the trike carriage assembly orthat are attached to the suspended trike in flight, including any win

18、g attachment bolts, shall be included in the trike carriageassembly empty weight, Wtkmt. These must include the required minimum equipment, unusable fuel, maximum oil, and whereappropriate, engine coolant and hydraulic fluid. Trike carriage empty weight, Wtkmt, shall be recorded in the Aircraft Oper

19、atingInstructions (AOI).3.1.4 wing weight, Wwing, nall parts, components, and assemblies that comprise the wing assembly, or that are attached to thewing in flight, shall be included in the wing weight, Wwing. The wing weight, Wwing, shall be entered in the AOI.3.2 Abbreviations:3.2.1 AOIAircraft Op

20、erating Instructions3.2.2 CCelsius3.2.3 CAScalibrated air speed (m/s, kts)3.2.4 cmcentimetre3.2.5 daNdeca Newton3.2.6 FFahrenheit3.2.7 Hgmercury3.2.8 IASindicated air speed (m/s, kts)3.2.9 in.inch3.2.10 ISAinternational standard atmosphere3.2.11 kgkilogram3.2.12 kt(s)nautical mile per hour (knot) (1

21、 nautical mph = (18523600) m/s)3.2.13 lbpound (1 lb = 0.4539 kg)3.2.14 mmetre3.2.15 mbmillibars3.2.16 NNewton3.2.17 psipounds per square inch gage pressure3.2.18 sseconds3.2.19 SIinternational system of units3.2.20 VAdesign maneuvering speed3.2.21 VCdesign cruising speed3.2.22 VDFdemonstrated flight

22、 diving speed3.2.23 VHmaximum sustainable speed in straight and level flight3.2.24 VNEnever exceed speed3.2.25 VS0stalling speed or minimum steady flight speed at which the aircraft is controllable in the landing configuration3.2.26 VS1stalling speed, or the minimum steady flight speed in a specific

23、 configuration3.2.27 Vxspeed for best angle of climb3.2.28 Vyspeed for best rate of climb3.2.29 VTmaximum aerotow speed3.2.30 WMAXmaximum design weight3.2.31 WSCweight shift control (aircraft)F2317/F2317M 1624. Flight Requirements4.1 Proof of Compliance:4.1.1 It shall be possible to demonstrate that

24、 the aircraft meets the requirements in this section at each allowable combinationof 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 theproduction aircraft except in the following case:4.1.2

25、.1 For the purposes of this test only, the aircraft may be modified to expand the control travel or limits in pitch whenestablishing VDF or VS1.4.1.3 Airspeeds shall be corrected to standard atmospheric conditions 1013.25 mb 29.92 in. Hg, 15C 59F.4.1.4 Climb performance requirements shall be met at

26、standard conditions or conditions more adverse.4.2 General Performance:4.2.1 Stall Speed in the Landing Configuration (VS0):4.2.1.1 The stall speed, if obtainable, or the minimum flight speed shall be established with: (1) engine idling with the throttleclosed, (2) hang point that produces the highe

27、st stalling or minimum flight speed, (3) maximum takeoff weight, and (4) trim settingin the landing configuration.4.2.1.2 VS0 shall be determined by flight-testing, in accordance with the following procedures: (1) aircraft power at idle, at aspeed of not less than VS0 plus 2.6 m/s 5 kts, and (2) the

28、 speed reduced at a rate not exceeding 0.5 m/s 1 kt/s until the stall isproduced 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

29、 and the recovery,or, if stall is not achieved before the control limit is reached, during the slowing to VS0 and subsequent acceleration to VS0 plus2.6 m/s 5 kts.4.2.2 Stall Speed Free of Control Limits (VS1):4.2.2.1 Where control limits result in VS0 being reached before the aircraft stalling, the

30、n the stall speed free of control limits(VS1) shall be determined. VS1 shall be established with: (1) the aircraft in the landing configuration defined in 4.2.1.1, and (2) theaircraft may be modified for the purposes of this test, only to expand the nose up pitch control range to the extent necessar

31、y forthe aircraft to stall when flown in accordance with the procedures detailed in 4.2.1.2.4.2.2.2 Where VS0 as determined in accordance with the procedures of 4.2.1.2 is the speed at which the aircraft stalls, then VS1= VS0.4.2.3 Minimum Climb Performance:4.2.3.1 The gradient of climb at recommend

32、ed 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 VibrationFlight-testing shall not reveal, by pilot observations, potentially damaging buffeting,airframe, or controls vibratio

33、n, flutter (with attempts to induce it), or control divergence, at any speed from VS0 to VDF.4.2.5 Turning Flight and StallsStalls 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

34、 limit of pitch control is reached. After the turning stallor reaching the limit of pitch control, level flight shall be regained without exceeding 60 of roll. This shall be performed with theengine at idle. No loss of altitude greater than 152 m 500 ft, uncontrolled turn of more than one revolution

35、, or speed buildup togreater than VNE shall be associated with the recovery.4.2.6 VHMaximum sustainable speed in straight and level flight, knots CAS.4.2.6.1 VH shall be established in straight and level flight with: (1) maximum allowed continuous engine power, and (2) thecombination of weight, load

36、ing, trimmer setting, and use of the flight controls allowed by the manufacturer that yields the highestsustainable speed.NOTE 2In the case where maximum continuous engine power results in a climb at maximum speed, power may be reduced as needed to maintainlevel flight.4.3 Controllability and Maneuv

37、erability:4.3.1 GeneralWhen operating in accordance with the recommendations in the Aircraft Operating Instructions, the aircraftshall 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

38、and off,4.3.1.6 With sudden engine failure,4.3.1.7 Turns,4.3.1.8 Changing speeds between VS0 and VNE, and4.3.1.9 Dive to VNE.4.3.2 Longitudinal Control:F2317/F2317M 1634.3.2.1 Starting at a speed of 1.1 VS0, it shall be possible to pitch the nose downwards so that a speed equal to 1.3 VS0 can bereac

39、hed in less than 4 s.4.3.2.2 It shall be possible to pitch the nose up at VNE at the most adverse hang point, trimmer setting, and engine power.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 theopposit

40、e direction. This shall be possible in both directions within 5 s from initiation of roll reversal, with the aircraft flown at 1.3VS0.4.3.3.2 Lateral control forces shall not reverse with increased displacement of the flight controls.4.3.4 Trim SpeedsThe speeds to achieve longitudinal trim shall lie

41、 between 1.3 VS0 and 0.909 VNE at all engine powers andthe allowable hang points.4.3.5 Ground HandlingIt shall be possible to prevent ground looping, with normal use of controls, up the maximum crosswindcomponent published in the AOI.4.4 Stability:4.4.1 Longitudinal Stability:4.4.1.1 The aircraft sh

42、all 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 attainand maintain any speed below trim.As the control force is reduced,

43、the aircraft shall return to within 20 % the original trim speed.4.4.2 Pitch TestingA test of the wing pitching moment about the hang point shall be conducted at VS0 0.866 over the rangeof angles of attack from 15 above zero lift angle to 10 below zero lift angle of attack. The wing shall exhibit a

44、trim angle abovezero lift angle of attack, and a positive pitching moment at any angle below trim, or if trim is not achieved in the test range, thewing shall exhibit a positive pitching moment throughout the range of angles specified.NOTE 3This test may be conducted as a taxi test with the wing mou

45、nted to the trike carriage.5. Structural Requirements5.1 Strength Requirements:5.1.1 Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimateloads (limit loads multiplied by prescribed factors of safety as specified in 5.3). Unless otherw

46、ise provided, prescribed loads arelimit loads.5.1.2 The structure shall be able to support limit loads without permanent deformation. At any load up to limit loads, thedeformation may not interfere with safe operation.5.1.2.1 The structure shall be able to support ultimate loads with a positive marg

47、in of safety (analysis) or without failure forat 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

48、requirements only if the structure conforms to those for whichexperience has shown this method to be reliable. Aerodynamic data required for the establishment of the loading conditions shallbe verified by tests, calculations, or conservative estimation.5.2.1.1 For analysis and test purposes, unless

49、otherwise provided, the air and ground loads shall be placed in equilibrium withinertia forces, considering each major item of mass in the aircraft. The loads shall be distributed so as to represent actual conditionsor a conservative approximation to them.5.2.2 If deflections under load would significantly change the distribution or amount of external or internal loads, thisredistribution shall be taken into account.5.2.3 The results obtained from strength tests should be corrected for departures fr