(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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(1) Structural substantiation is required for the gear, retracting mechanism, doors, gear supporting structure for landing loads, maneuvering, gusts, and yawing flight condition loads.
(2) An emergency means to extend the gear after failure of the retraction/extension system is required for all except solely manual mechanical systems.
(3) This regulation requires an indication to the pilot when the gear is secured in the extreme positions. This rule does not apply to rotorcraft with fixed gear. The rule also applies to amphibious rotorcraft with retractable gear.
(4) A landing gear down-lock is required. An optional up- lock may be used if it meets reliability requirements.
(5) A (ground) operation test must be conducted to ensure proper functioning of the system.
(6) Location and operation of the control lever or device must comply with § 29.777. This section includes identification of controls to prevent confusion and inadvertent operation. Sections 25.779 and 25.781 of FAR Part 25 contain large airplane design requirements for motion, effect, and shape of cockpit controls and their knobs and should be consulted for further guidance.
b. Procedures.
(1) The design load factors and resulting loads should be derived from the design data. The landing gear, while retracted, operating, and extended, and its supporting structure should be substantiated for the critical aerodynamic and inertia loads. Yawed conditions should be considered. The specific conditions are noted in paragraphs (a)(1), (2), and (3) of § 29.729.
(2) Wheel well doors, if installed, should be designed for the aerodynamic loads, including loads from yawing conditions (angles proven under § 29.351) for airspeeds up to the design maximum landing gear extended speed. Aerodynamic effects on both open and closed doors must be considered in the door and door support substantiations. The applicant may choose to substantiate the rotorcraft for a “landing gear operating” and “extended” speed VLO and VLE, respectively, that is equal to the rotorcraft VNE. This option will alleviate an airspeed “structural limitation” because of the landing gear design substantiation. Any airspeed “structural limitation” should be listed in the structural limitations part of the TIA.
(3) The required “down-lock” should be checked during the operation test. The design drawing should be reviewed for compliance prior to conducting an operation test.
The “down-lock” system should be evaluated for § 29.1309 function and reliability requirements.
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(4) If an optional “up-lock” is installed (including hydraulic locking), the landing gear should be extended during the operation test after simulation of critical failure mode of the retraction system (reference § 29.1309).
(5) An “operation” test plan or proposal submitted for compliance with § 29.729(d) should include the items noted in the two previous subparagraphs and should include a functional check of the position indicator system. Those tests must be satisfactorily completed before issuing the TIA.
(6) During the official FAA/AUTHORITY flight tests, compliance with the emergency operation, position indicator, and control aspect of § 29.729(c), (e), and (f), respectively, will be verified or accomplished. In addition, the F and R test program plan (§ 21.35) will specify certain tests or evaluations for the retraction system.
(7) Position Indicator Evaluation.
(i) When evaluating the position indicator system, emphasis should be placed on the switches and their installations, and on the cockpit presentation. Each gear must have its own set of switches to indicate when it is secured in its extreme “up” position and its extreme “down” position. The switches must be located to give a valid indication of the arrival of the gear at its extreme position.
(ii) The reliability and environmental qualifications of the switches to be used should be carefully considered. An example of a condition that has potential for trouble is operation on wet areas. Trouble starts when water is picked up by the tires and deposited on the switches. During winter months the water can freeze, and the resulting ice may prevent the switch from functioning properly.
(iii) An acceptable cockpit presentation consists of two lights for each gear. One light is colored “green” and indicates when its gear is secured in the extreme “down” position. The other light is colored “amber” or “red” and indicates when its gear is in transit. When the gear is in either extreme position, the in transit light is “out.” For this presentation, the indication to the pilot that the gear is in the extreme “up” position is an all-gear, lights-out condition.
(iv) Some manufacturers have also included a warning system to alert the crew if the landing gear has not been extended prior to landing. If a warning system is presented, §§ 29.1301 and 29.1309 should be used to evaluate its functional characteristics and the impact of its failure modes.
AC 29.731. § 29.731 WHEELS.
a. Explanation. This standard requires use of approved wheels, either approved under TSO-C26 or a later revision or approved under the type certificate for the aircraft.
Wheels must satisfy both a design static (1g) load and design limit landing or taxiing Page D - 55
load determined under the applicable ground load requirements. Standards for a tire installed on a wheel are contained in § 29.733.
b. Procedures.
(1) The structural design loads data shall contain both a static load and a landing and taxiing load for each wheel. These loads are determined by virtue of compliance with the standards of § 29.731(b) and (c). The ratings of the wheel shall not be exceeded. TSO -C26c contains minimum performance standards for TSO approval of aircraft wheels and wheel-brake assemblies. Ratings are assigned in accordance with this performance standard.
(2) If a wheel selected for an aircraft design has TSO approval, the wheel manufacturer will supply the rating to the aircraft manufacturer. Each wheel shall be marked as prescribed which includes a listing of the TSO number. Even though a wheel is TSO approved, the application on the aircraft (loads imposed on the wheel) requires proof that the rating is not exceeded.
(3) If a wheel selected for an aircraft design is not approved under a TSO, the necessary data, both detail design and assembly drawings and qualification tests and test report data, will be required to comply with the standards contained in Part 29.
Design control and inspections will be accomplished as a part of the aircraft type design. Structural substantiation and any appropriate qualification tests shall be accomplished. See §§ 29.471 through 29.497 and § 29.511 for the ground load conditions.
(4) The Tire and Rim Association, Inc., generally issues a yearbook listing tire and rim sizes and ratings. The dimensions and contours for aircraft wheel rims are contained in Section 9 of this yearbook.
AC 29.733. § 29.733 (Amendment 29-12) TIRES.
a. Explanation.
(1) This standard specifies both design and performance criteria for tires. The tire must fit the wheel rim. The maximum static ground reaction for the condition specified must not exceed the maximum static load rating of each tire. In addition, any tire of retractable gear systems must have adequate clearance from surrounding structure and systems as specified.
(2) Main, nose, and tail wheel tires must comply.
(3) Rotorcraft design maximum weight shall be used. Static and “dynamic” conditions are specified for rotorcraft tires.
(4) Tire performance standards are contained in TSO-C62c.
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b. Procedures.
(1) The aircraft structural design loads should contain a maximum static load imposed on the tires. The load is derived for a static ground reaction assuming the design (maximum) weight and the critical center of gravity for each tire of the landing gear. The wheel loads are determined under § 29.731(b). Reduced weight but forward CG conditions may result in the highest static load on a nose wheel tire. Thus, combinations of weight and CG locations require investigation for the maximum tire load of each main, nose, and tail wheel tire. Nose wheel tires are subject to a specific dynamic condition.
(2) The maximum possible size of the tires considering appropriate temperatures, aging, and pressure should be obtained to check wheel well and cover clearances. Tire dimensions (for clearances) may be found in the yearbook noted in paragraph AC 29.733b(4). If the tire clearance is questionable, objects may be taped to the tire to simulate tire growth or oversize dimensions expected and the wheel retracted and rotated by hand to check for possible interferences. Minimum clearance, such as one-half inch, may be adequate as a design objective. The design drawings should be reviewed for information of correct systems installations and landing gear rigging within the wheel wells and wheel covers, if installed. If necessary to control tire sizes, specific manufacturer’s tires should be used as “required equipment” and the tire manufacturer and the part number should be specified in the design data and on the type certificate data sheet as “required equipment.”
(3) As specified in § 29.729(d), an operation test of any retractable landing gear should be performed. During this operation test, the tire clearances should be determined and recorded for the maximum tire size expected in service. Only the least or minimal clearance found, if adequate, should be recorded.
(4) The Tire and Rim Association, Inc., generally issues a yearbook listing tire and wheel rim sizes and ratings. This information is advisory as stated in the yearbook.
Section 9 concerns aircraft tires and rims. Table AP-5 in Section 9 of the yearbook concerns tires used on rotorcraft. The tire may be selected initially from the yearbook, but qualification data for the specific tires used shall be furnished with the type design data in compliance with the standards. Section 9 also contains tire size and tire growth dimensions.
(5) Minimum performance standards for aircraft tires, excluding tail wheel tires are found in TSO-C62c, Aircraft Tires. Tires meeting the TSO are marked as prescribed in the standards. The load rating (reference § 29.733) is marked on the tire.
TSO tires are not required but should be used whenever possible. The manufacturer’s information, such as load rating, should be included in the aircraft type design structural substantiation data.
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