AC 29-2C, Chg 3 9/30/2008 AC 29.681. § 29.681 LIMIT LOAD STATIC TESTS.
a. Explanation.
(1) The rule requires static tests of the control system in showing compliance with limit load requirements.
(2) The tests are specified to include each fitting, pulley, and bracket of the control system being tested and to include the “most severe loading.”
(3) Also, the rule requires that compliance with bearing factors (reference § 29.623) be shown by individual tests or by analyses for control system joints subject to motion.
b. Procedures.
(1) Compliance with the requirements of this rule is obtained by static tests conducted on either a static test airframe or on a prototype flying ship. In either case, conformity of the control system and related airframe is necessary to validate the tests.
(2) The rotor blades or aerodynamic surfaces may be used to react pilot effort loads through the control system or they may be replaced with fixtures. If fixtures are used, they should be evaluated for geometric and stiffness effects to assure test validity.
(3) The loads to be applied during the limit load static tests are specified in §§ 29.395, 29.397, and 29.399. The loads are applicable to collective, cyclic, yaw, and rotor blade control systems as well as any other flight control systems provided by the design.
(4) Section 29.585(e) specifies bearing factors for control system joints subject to angular motion. These factors are 3.33 for push- pull systems and 2.0 for cable systems for joints with plain bearings. For joints with ball or roller bearings, use the manufacturer’s ratings.
AC 29.683. § 29.683 OPERATION TESTS.
a. Explanation. The rule requires that the control system be free from jamming, excessive friction, and excessive deflection. An operational test is required in which specified loads are applied at the pilot controls and carried through an operating control system.
b. Procedures.
(1) Compliance with the requirements of this rule is obtained by use of a test setup similar to that used for the limit load tests of § 29.681, except the load reactions at Page D - 42
the blades (or surfaces) must allow for movement of the blades (or surfaces) as the system is operated through its operating range.
(2) Fixtures are normally affixed to the surfaces (or replace the surfaces) to allow pulley arrangements which provide for movement under load. These fixtures should be evaluated to assure that system loads up to limit will be applied during the full range of operations of each system.
(3) Each flight control system should be operated through its entire range under a light load and under limit load. As the controls are being operated, the system should be checked for jamming, excessive friction, and excessive deflection. Excessive deflection includes deflection sufficient to contact other systems or structure. Also, if under these limit load conditions the components deflect, the deflection would be considered excessive if there is permanent deformation of any component or supporting structure. Also any deflection that results in an uncorrected condition when the load is released, e.g., if a bellcrank is forced off-center or over-center during load and does not return to the normal position after load release is excessive deflection. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system. However, care should be taken when removing panels so that airframe structure is not weakened enough to deflect from its normal position when test loads are applied to the control system.
AC 29.685. § 29.685 (Amendment 29-12) CONTROL SYSTEM DETAILS.
a. Explanation. The rule requires that the control system be designed to prevent chafing, jamming, and interference from cargo, passengers, loose objects, or the freezing of moisture. Specifically, means are required in the cockpit to prevent the entry of foreign objects into places where they would jam the system, and means are required to prevent the slapping of cables or tubes against other parts. Specific design considerations to prevent binding and overloads within the control system are required such as--
(1) Assure pulley-cable combinations as specified in MIL-HDBK-5 are used unless inapplicable.
(2) Assure close fitting pulley guards are provided.
(3) Assure pulley-cable alignment sufficient to prevent excessive pulley flange loads is provided.
(4) Assure fairlead-cable alignment is within 3°.
(5) Assure no clevis pins are retained only by cotter pins.
(6) Assure turnbuckles do not bind other structures throughout the range of travel.
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AC 29-2C, Chg 3 9/30/2008
(7) Assure means for inspection of control system components are provided.
(8) Assure control system joints subject to angular motion incorporate special bearing factors, 3.33 for push-pull systems and 2.0 for cable systems.
(9) Assure that manufacturer’s ratings for ball or roller bearing ratings are not exceeded.
b. Procedures.
(1) The geometry of the control system components and installations is the primary control to prevent chafing, jamming, and interference. The control system from cockpit to surface should be checked for clearances both unloaded and loaded. The control system should be checked under load during both the limit load static tests (reference § 29.681) and the operational tests of § 29.683. Location of guides or fairleads and pulleys may be used in cable systems to prevent chafing and interference with other structure. Generally, tubes should clear adjacent structure by location and design geometrical considerations. If supplemental means are provided to assure the tubes do not chafe or interfere, the means should be evaluated for possible jamming.
(2) Rubber (or other elastomeric) boots connected to both the cockpit control arm or shaft and to the floor are acceptable means to prevent the entry of foreign objects into underfloor areas where they may cause jamming of controls. Control systems should, in general, be routed around cargo compartments. If routing of the control system components is in or near cargo areas, the control system components should be protected by bulkheads, panels, or other enclosures which have sufficient strength and stiffness to prevent possible interference with the control system components when subjected to cargo loading and handling deflections.
(3) Control system details should be reviewed for possible moisture collection.
Areas should drain free. Exposed or open control areas should drain free, and areas of possible freezing moisture collection should not accumulate ice that would cause a jam of the controls. Simulated or actual ice collection on the controls may be used to prove questionable features. The areas to be considered for moisture collection include both external and internal areas where moisture may accumulate by direct impingement of water, entrapment of water particles, or condensation of moisture.
(4) The latest revisions of MIL-HDBK-5 do not explicitly give approved pulley-cable combinations, but appropriate MIL specifications are given in Chapter 8.3 for use in determining pulley-cable combinations and ratings.
(5) Provide ratings, factors, and alignment as specified.
(6) Provide inspection means as specified.
(7) Provide close fitting pulley guards as specified.
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AC 29.687. § 29.687 SPRING DEVICES.
a. Explanation.
(1) This standard for control systems assures that springs and spring devices used to prevent flutter, control oscillations, or vibrations are either --
(i) Reliable; or
(ii) The failure is not critical to the rotorcraft.
(2) Tests simulating service conditions are required in either instance.
b. Procedures.
(1) Springs and spring devices used in the control system, including balance springs, should be identified early in the certification program.
(2) If a spring cannot be shown by observation or analysis to be noncritical, then ground or flight tests may be required.
(3) Springs that are critical to safe operation may be subject to fatigue substantiation to prove they are reliable for the operating conditions imposed in service.
(4) Springs used in conjunction with hydraulic actuator spool valves may be subject to the standards of § 29.695.
AC 29.691. § 29.691 AUTOROTATION CONTROL MECHANISM.
a. Explanation.
(1) Rotorcraft designs generally have a main rotor blade collective pitch control system that does not have detents or other devices to limit pitch control in the control mid-range. Autogyro and other rotorcraft designs may include detents or other finite position control for collective pitch control. This rule requires that the control design allow rapid entry into autorotation after a power failure.
(2) Section 29.33 contains standards concerning establishment and control of the main rotor speed limits. The standard requires flight tests and demonstrations. The standard also concerns rotorcraft design features that are related to control of the main rotor speed limits. Paragraph AC 29.33, § 29.33, pertains to this standard.
(3) Other design requirements for control systems are contained in § 29.685.
b. Procedures.
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(1) If high and low main rotor pitch stops are employed in the collective control and if the control may be rapidly moved from one limit to the other, compliance is shown.
(2) If detents or intermediate stops are employed, the pilot must be able to easily and readily override, disconnect, remove, or bypass the device to allow rapid autorotation entry prior to exceeding transient low speed rotor limits. An early assessment for design deficiencies may be accomplished by the flight test personnel with the evaluation completed in the Type Inspection Authorization (TIA) test program.
(3) It is acknowledged that modern rotorcraft designs may have an autorotation VNE that is lower than power-on VNE or normal cruise speed. For rotorcraft designs with this characteristic, the speed must be reduced after entry into autorotation. The rule also applies to rotorcraft designs with this characteristic and no relief from the rule is required since many phases of operation occur at speeds less than power-on VNE. For example, a critical phase of flight occurs during takeoff. Rapid entry into autorotation is essential during this phase also.
(4) The features of the autorotation control mechanism and ability to control the rotor speed within the design limits for any rotorcraft will be evaluated as an integral part of the TIA test program.
AC 29.695. § 29.695 POWER BOOST AND POWER-OPERATED CONTROL SYSTEM.
a. Reference Regulations. The following sections of Part 29 are either incorporated in the provisions of § 29.695 or are otherwise applicable to power boost and power-operated control systems:
(1) Section 29.307Proof of structure.
(2) Section 29.571Fatigue evaluation of fight structure.
(3) Section 29.681Limit load static tests.
(4) Section 29.685Control system details.
(5) Section 29.861Fire protection of structure, controls, and other parts.
(6) Section 29.863Flammable fluid fire protection.
(7) Section 29.1301 Function and installation.
(8) Section 29.1309 Equipment, systems, and installations.
b. Explanation.
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(1) The rule requires an alternate system if a power boost or power-operated control system is used.
(2) The alternate system must, in the event of any single failure in the power portion of the system, or in the event of failure of all engines:
(i) Be immediately available.
(ii) Allow continued safe flight and landing.
(3) The alternate system may be:
(i) A duplicate power portion of the system; or
(ii) A manually operated mechanical system.
(4) The power portion of the system includes:
(i) The power source (such as hydraulic pumps); and
(ii) Items such as valves, lines, and actuator.
(5) The failure of mechanical parts (such as piston rods and links) must be considered unless their failure is extremely improbable.
(6) The jamming of power cylinders must be considered unless their jamming is considered extremely improbable.
c. Procedures. It is assumed in the following discussion that the power boost or power-operated control system being utilized is a typical aircraft hydraulic system.
(1) The rule requires, without regard to the probability of failure, an alternate system for the power portion of the system. The power portion of the system, by example in the rule, includes hydraulic pumps, valves, lines, and actuators. It has also been interpreted to include seals, servo valves, and fittings.
(2) If a duplicate power portion of the system is used to meet the requirements of the rule, the requirements may be met by providing a dual independent hydraulic system, including the reservoirs, hydraulic pumps, regulators, connecting tubing, hoses, servo valves, servo-valve cylinder, and power actuator housings. There must be no commonality in fluid-carrying components. A break in one system should not result in fluid loss in the remaining system.
(3) Dual actuators should be designed to assure that any single failure in the duplicated portion of the system, such as a cracked housing, broken interconnecting Page D - 47
AC 29-2C, Chg 3 9/30/2008 input, or broken interconnecting output link, does not result in loss of total hydraulic system function.
(4) A manually operated mechanical system may be used as the alternate system to a single hydraulic system if, after the loss of the single hydraulic system, the pilot can control the rotorcraft without exceptional piloting skill and strength in any normal maneuver for a period of time as long as that required to effect a safe landing.
The control forces should not exceed those specified in § 29.397 and flight characteristics should meet the requirements of §§ 29.141 (b) and (b)(3).
(5) The substantiation of the various system components should include consideration for operation in the normal and alternate system modes.
(6) The “extremely improbable” criteria noted in § 29.695(c) for failure of mechanical parts may be satisfied by performing component fatigue testing and establishing a service life through this technique.
(7) Fatigue substantiation of the control actuator is required under § 29.571 and should consider both the stresses imposed by flight loads and the stresses imposed by hydraulic pump pressure pulses. Flight loads factored in a suitably conservative manner may be an acceptable means to take into account both effects.
(8) The possibility of jamming of the power cylinder may be shown as “extremely remote” through a failure analysis that considers every possible system component failure such as, but not limited to, ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, jammed servo valves, broken interconnecting servo valve inputs, broken interconnecting output links, etc.
(9) Three acceptable means to meet the requirements of § 29.695(a)(2) could be as follows:
(i) Provide two transmission-driven hydraulic pumps, provided the pumps are driven by the transmission during all flight conditions including autorotation.
(ii) Use two electrically driven hydraulic pumps if electrical power is available to drive the pumps with all engines failed. If this approach is used, the battery must be capable of running both pumps plus all other required equipment necessary for continued safe flight.
(iii) Use a single transmission driven pump and an electrically driven pump.
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SUBPART D - DESIGN AND CONSTRUCTION
LANDING GEAR AC 29.723. § 29.723 SHOCK ABSORPTION TESTS.
a. Explanation.
(1) Limit and “reserve energy” drop tests are required as prescribed in §§ 29.725 and 29.727, respectively. These tests must be conducted on the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation. For rotorcraft with skid landing gear, the tests may be conducted on the complete rotorcraft or on a simulated fuselage with the complete skid landing gear system.
(2) The rotorcraft must be designed to limit load factors that equal or exceed the limit load factor substantiated by these drop tests. In practical application, the rotorcraft may be designed to a limit load factor, such as 2.8 g. Thus, it is necessary that the limit landing load factor derived from the landing gear drop tests be equal to or less than 2.8 g. If not, the rotorcraft must be redesigned for the higher load factor derived from the drop tests. It must be shown in accordance with § 29.723 that the limit load factors selected for design under § 29.473 will not be exceeded in landings with the limit descent velocity corresponding to the drop height specified in that section. In addition, reserve energy absorption capacity of the landing gear must be shown for a descent velocity of 1.22 times the limit descent velocity selected under § 29.473 by increasing the drop height to 1.5 times the “limit” drop height. The test requirements or procedures outlined in FAR 29 for obtaining the landing load factors are empirical;
however, these procedures are based on and supported by satisfactory experience.
(3) As stated in § 29.725(c), each landing gear unit should be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it.
(i) For wheel landing gear designs, the level landing or tail down landing and level landing with drag are generally the most critical attitude. A test of more than one attitude is generally required to comply with the standard. The landing attitudes or conditions prescribed are level (vertical loads), inclined (loads at 14.5° aft from the vertical axis), level with wheel spin-up and tail down. These attitudes are specified in §§ 29.479(b)(1), (2), and (3) and 29.481.
(ii) For skid landing gear designs, the level landing and level landing with drag are generally the most critical attitudes. These attitudes are specified in § 29.501(b) and (c).
(4) Drop tests are required. If analytical methods and/or means are proposed by the applicant, the data presented for approval must be equal to or conservative with Page D - 49
respect to that data obtained from physical drop tests. Section 21.21(b)(1) of FAR 21 concerns “equivalency” determinations. Presenting an acceptable means of “equivalency” here would circumvent the necessary scrutiny of an analytical method or means and is also beyond the scope of this document.
b. Procedures. The test plan or proposal must be approved prior to official FAA/AUTHORITY tests unless satisfactory resolution of outstanding proposal or conformity inspection items can be accomplished after the test.
(1) The following headings would be a typical table of contents for the test proposal, and a generalized explanation of the contents that may be included under each of these headings for a wheel landing gear follows.
(i) Purpose. The regulations to which compliance is being shown by the drop tests should be identified (usually §§ 29.723, 29.725, and 29.727). Also, the rotorcraft landing gear including the wheels and tires to be dropped, should be positively identified in the report by the manufacturer’s or applicant’s previously FAA/AUTHORITY approved drawing, technical standard orders (TSO’s), or other identifying FAA/AUTHORITY approved data as applicable.
(ii) Description of test setup. This section should present a description of the test fuselage or jig, method of attaching landing gear to jig, and type of accelerometer to be used to measure load factors. Proof of calibration of accelerometer should be available. The accelerometer should be mounted at the aircraft CG if a free drop of the aircraft is used, or as close as practical to the centerline of the main shock absorbing component of each landing gear (oleo strut, etc.) if each gear is tested separately. The description of the test jig, including platforms on which the gears are to be dropped, should be defined by sketches in addition to the required mathematical calculations. This data should show that the landing gear will be at the proper attitude, relative to the platform, on impact for the particular landing condition. Drawings or other approved data from which the geometry is taken should be referenced in the proposal.
The tire and oleo pressures at the time of the test should be specified. The method of measuring the deflection of the tire plus the vertical travel of the axle under impact should be described. This measurement may be accomplished by telescoping tubes attached to the point on the jig that would measure the total (tire and oleo) vertical deflection of the landing gear. Other vertical and horizontal deflections should be measured as required to determine if the landing gear has experienced permanent deformation after each drop test. The effect of surface roughness should be considered. Smooth surfaces tend to give maximum deflections where rough surfaces tend to restrict deflection and to result in maximum values of NZ. Preliminary company drop tests (at less than limit drop height) may be used to determine the critical surface roughness, or engineering evaluations may be used (without tests) when the gear configurations are such that the critical surface condition can be analytically determined (or when the load factor is shown to be negligibly affected by surface roughness).
NACA Report 1154, dated 1953, contains information that surface coefficients of friction may vary from 0.4 to 0.7. Skid landing gear standards, § 29.501(c), indicate an Page D - 50
acceptable coefficient of friction is 0.5. A wheel landing gear design standard, § 29.479(b), indicates an acceptable coefficient of friction is 0.25. In the case of a small rotorcraft, the entire aircraft may be dropped. This may be accomplished by establishing pivot points at the main gear axles for the tail (or a point forward of the nose gear) drops, and a pivot point at the tail (or nose gear) axle for the main gear drops. It is the responsibility of the applicant to distribute the aircraft inertia items, including added weight to get the proper effective drop weight (We) at the landing gear, so that no local failures of the aircraft occur as a result of the limit or reserve energy drop tests.
(iii) Test data. Computations for the required drop height (h) and the effective drop weight (We) should be shown for each design level landing and tail down landing condition in compliance with §§ 29.479 and 29.481. The computations should be in accordance with § 29.725(a) for h and § 29.725(b) for We for the limit drop tests.
We and h are computed in accordance with § 29.725 for the limit drop test and with § 29.727 for reserve energy drop test. The computation of the static weight on the gear being dropped (WM, WN, or WT) and used in the computation of We should be shown.
This static weight is defined as WM, WT, or WN for the main gears, tail gear, or nose gear, respectively, in § 29.725(d). It should be shown that the critical CG and proposed certificated maximum landing weight have been used in the computation of WM, WT, or WN. The computation of the slope of the platforms required for the inclined reacti on conditions should be presented also.
NOTE: Effective drop weight (W E) is used only for free drops. It provides a technique for accounting for rotor lift without applying lift during the test. If rotor lift is applied during the drop tests, actual weights (WM, WT, or WN) will be used, not effective weights, WE.
(iv) Test results. The results of the test are based on the values of W E, h, d, W, and L used and obtained for each drop test and the value of Nj obtained from the accelerometer. These results should be summarized, and the method of computing the aircraft limit inertia load factor should be shown for each drop in accordance with § 29.725(d). A print or copy of the film or other recording trace from the accelerometer, if not a direct readout type of accelerometer, should be included in the test results.
Each critical condition should have several preliminary drops as many times as required to obtain reasonable correlation.
(2) Skid landing gear may be tested using similar procedures except a level landing attitude drop test is all that is required by § 29.501. The design load conditions specified in § 29.501(c) through 29.501(f) are derived from this level drop test condition.
(i) Section 29.501, paragraphs (a)(2) and (3), contain special considerations for skid landing gear.
(ii) Section 29.501(a)(2) specifies that structural yielding of elastic spring members under limit load is acceptable. This yielding or deformation is a means of Page D - 51
absorbing the landing impact. For skid landing gear that use oleo or other types of shock absorbers, the standard does not allow structural yielding under limit load.
During the limit load and reserve energy (ultimate for skid landing gear with elastic spring numbers) drops, the yielding energy absorbing members will probably deform or yield. After a limit drop test, the gear may be used for a reserve energy drop at the discretion of the applicant but a gear that has been subjected to a reserve energy drop should not be used unless it can be shown that no yielding has occurred in that gear.
(3) Wheel landing gear is tested in attitudes prescribed in paragraph a(3)(i) above. Each unit, nose or main gear, is generally tested separately.
(4) Skid landing gear is tested in attitudes prescribed in subparagraph a(3)(ii) above. Due to the construction of skid landing gear, the complete skid landing gear is tested as a unit. Thus, the level landing with drag condition is probably the critical attitude for the forward cross-tube and its attachments. The level landing condition is probably the critical attitude for the aft cross-tube and its attachments.
(5) An FAA/AUTHORITY or FAA/AUTHORITY designated or delegated person need only witness the drop tests for “record” or “compliance.” Preliminary or developmental drops do not require an FAA/AUTHORITY witness.
AC 29.725. § 29.725 (Amendment 29-3) LIMIT DROP TEST.
a. Explanation. Limit drop tests, in the critical aircraft attitude or critical attitude of each gear, are required for the landing gear. The drop height must be at least 8 inches, which equates to 393 feet per minute (free fall) vertical descent speed. Rotor lift may be simulated and an effective mass may be used in the drop test as prescribed.
b. Procedures. See paragraph AC 29.723, § 29.723.
AC 29.727. § 29.727 RESERVE ENERGY ABSORPTION DROP TEST.
a. Explanation.
(1) In addition to the limit drop tests, a reserve energy drop test is required.
The landing gear must not collapse in this test to the extent that the fuselage impacts the ground. Fracture (to separation) of landing gear parts is considered collapse of the landing gear. This test is not an ultimate load drop test for the landing gear, except as specified in § 29.501(a)(3) for certain skid landing gear designs using elastic spring members.
(2) All other types of landing gear must be substantiated for design ultimate loads in addition to this reserve energy drop test.
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(3) Shock absorbing devices, such as oleos, must not “bottom” during the reserve energy drop test. “Bottoming” occurs when displacement of the device no longer occurs with increasing load.
(4) Requirements for proof of the landing gear and airframe structure are found in §§ 29.305, 29.307, and 29.473.
b. Procedures. See paragraph AC 29.723, § 29.723.
AC 29.727A. § 29.727 (Amendment 29-30) RESERVE ENERGY ABSORPTION DROP TEST.
a. Explanation. Amendment 29- 30 defines the word “collapse” as used in § 29.727(c). Collapse of the landing gear during reserve energy absorption drop tests occurs when:
(1) A member of the landing gear will not support the rotorcraft in the proper attitude; or,
(2) A member deforms sufficiently to allow the rotorcraft structure other than the landing gear and external accessories to impact the landing surface.
b. Procedures. The procedures of paragraph AC 29.727 continue to apply with the following supplemental guidance.
(1) The proper attitude for the rotorcraft after the reserve energy absorption drop test is an attitude which allows for permanent deformation of landing gear elements but provides for adequate egress from the rotorcraft.
(2) External accessories that may not impact the landing surface during drop testing include devices such as externally mounted fuel tanks or accessories likely to cause post-landing fires. Cameras, loudspeakers, and search lights may be damaged during deformations resulting from reserve energy drop tests if electrical connections are sufficiently protected to preclude electrical fires and the devices are not likely to penetrate fuel tanks. The expendable accessories, if installed, should also be designed to not have “hard points” that would unacceptably damage the rotorcraft structure under landing impacts by penetration into the occupied areas or fuel tanks. These expendable accessories should be designed with frangible fittings, frangible devices, or comparable design features. Also, these devices should be designed to not significantly alter the energy absorbing ability or design features of the landing gear.
AC 29.729. § 29.729 (Amendment 29-24) RETRACTING MECHANISM, LANDING GEAR.
a. Explanation .
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