287. § 27.687 SPRING DEVICES.
a. Explanation.
(1) This standard for control systems ensures that springs and spring devices used to prevent flutter, control oscillations, or vibrations are either -- (i> Reliable (failure is extremely remote); 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) Whenever a spring cannot be proven by observation or analysis that it is "not critical," 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 § 27.695.
288. § 27.691 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 midrange. Autogyro and other rotorcraft designs may include detents or other finite position control for collective pitch control. This rule requires that the control design allows rapid entry into autorotation after a power failure.
(2) Section 27.33 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.
(3) Other design requirements for control systems are contained in § 27.685.
b. Procedures.
Par 287 355
(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 autorotational entry prior to exceeding transient low speed rotor limits. An early assessment of the design 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 autorotational 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. 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 autorotational 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.
289. § 27.695 POWER BOOST AND POWER-OPERATED CONTROL SYSTEM.
a. Reference Regulations. The following sections of Part 27 are either incorporated in the provisions of § 27.695 or are otherwise applicable to power boost and power-operated control systems:
(1) Section 27.307 Proof of structure.
(2) Section 27.571 Fatigue evaluation of fight structure.
(3) Section 27.671 Control system.
(4) Section 27.681 Limit load static tests.
(5) Section 27.687 Spring devices.
(6) Section 27.685 Control system details.
(7) Section 27.861 Fire protection of structure controls and other parts.
(8) Section 27.863 Flammable fluid fire protection.
356 Par 288
(9) Section 27.1301 Function and installation.
(10) Section 27.1309 Equipment, systems, and installations.
b. Explanation.
(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
(il) A manually operated mechanical system.
(4) The power portion of the system includes:
The power source (such as hydraulic pumps); and
(il) 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 respect 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, Par 289 357
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 ensure that any single failure in the duplicated portion of the system, such as a cracked housing, broken interconnecting 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 undue mental or physical fatigue in any normal maneuver for a period of time as long as that required to effect a safe landing.
(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 § 27.695(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 § 27.571 and should consider both the stresses imposed by flight loads and the stresses imposed by hydraulic pump pressure pulses. Flight loads factored in a conservative way 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 § 27.695(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 sate flight.
(iii) Use a single transmission-driven pump and an electrically driven pump.
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7/30/97 АС 27-1A
290.-296. RESERVED.
Intentionally Left Blank Par 290 359 (thru 365)
297. RESERVED.
Intentionally Left Blank Par 297 366
7/30/97 АС 27-1A SECTION 18. LANDING GEAR
298. § 27.723 SHOCK ABSORPTION TESTS.
a. Explanation.
(1) Limit and "reserve energy" drop tests are required as prescribed in §§ 27.725 and 27.727, respectively. These tests may 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 2.8g. Thus, it is necessary that the limit landing load factor derived from the landing gear drop tests be equal to or less than 2.8g. If not, the rotorcraft must be redesigned for the higher load factor derived from the drop tests. It must be shown in accordance with § 27.723 that the limit load factors selected for design under § 27.473 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 § 27.473 by increasing the drop height to 1.5 times the "limit" drop height. The test requirements or procedures outlined in Part 27 for obtaining the landing load factors are empirical;
however, these procedures are based on and supported by satisfactory experience.
(3) As stated in § 27.725(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. 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 may be required to comply with the standard.
(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 respect to that data obtained from physical drop tests. Section 21.21(b)(1) 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.
Par 298 367
(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.
Purpose. The regulations to which compliance is being shown by the drop tests should be identified (usually §§ 27.723, 27.725, and 27.727). 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.
(її) 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 surtace 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, § 27.501(c), indicate an acceptable coefficient of friction is 0.5. A wheel landing gear design standard, § 27.479(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 368 Par 298
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 §$ 27.479 and 27.481. The computations should be in accordance with § 27.725(a) for h and § 27.725(b) for We for the limit drop tests.
We and h are computed in accordance with § 27.725 for the limit drop test and with § 27.727 for reserve energy drop test. The computation of the static weight on the gear being dropped (W, WN, or WT) and used in the computation of We should be shown.
This static weight is defined as WM, WT, or Wy for the main gears, tail gear, or nose gear, respectively, in § 27.725(d). It should be shown that the critical CG and proposed certificated maximum landing weight have been used in the computation of Ww, WT, or WN. The computation of the slope of the platforms required for the inclined reaction conditions should be presented also.
(iv) Test results. The results of the test are based on the values of We, h, d, W, and L used and obtained for each drop test and the value of N; 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 § 27.725(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 § 27.501. The design load conditions specified in § 27.501(c) through (t) are derived from this level drop test condition.
Section 27.501(a)(2) and (3), contain special considerations for skid landing gear.
(її) Section 27.501(a)(2) specifies that structural yielding of elastic spring members under limit load is acceptable. This yielding or deformation is a means of absorbing the landing impact. For skid landing gear that uses 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 298a(3).
Each unit, nose or main gear, is generally tested separately.
Par 298 369
(4) Skid landing gear is tested in attitudes prescribed in Paragraph 298a(3).
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.
299. § 27.725 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 a 393-foot-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 298, § 27.723, of this advisory circular.
300. § 27.727 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 § 27.501(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.
(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 §§ 27.305, 27.307, and 27.473.
b. Procedures. See Paragraph 298, § 27.723, of this advisory circular.
370 Par 298
300A. § 27.727 (Amendment 27-26) RESERVE ENERGY ABSORPTION DROP TEST.
a. Explanation. Amendment 27-26 defines the word "collapse" as used in § 27.727(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 landing gear 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 300 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. Refer to Paragraph 340 of this AC for emergency exit standards that relate to attitudes after a crash, § 27.807(b)(2).
(2) External accessories that may not impact the level landing surface during drop testing (or equivalent gear deflections) include devices such as externally mounted fuel tanks or accessories likely to cause post-landing fires. Expendable accessories, such as cameras, loudspeakers, and search lights, may be damaged during landing gear 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 a fuel compartment or occupied areas. 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. Design features may be employed to preclude this penetration if possibly hazardous. The expendable accessories, if installed, 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.
301. § 27.729 (Amendment 27-21) RETRACTING MECHANISM.
a. Explanation.
(1) This standard was added by Amendment 27-21.
(2) Structural substantiation is required for the gear, retracting mechanism, doors, gear supporting structure for landing loads, maneuvering, gusts, and yawing Par 300A 371
АС 27-1A 7/30/97 flight condition loads. Design maximum airspeed for extension and retraction and for fully extended conditions are required conditions.
(3) An emergency means to extend the gear after failure of the retraction/extension system is required for all except solely manual mechanical systems.
(4) This regulation requires an indication to the pilot when the gear is secured in the extreme positions. This rule does not apply to rotorcraft that have fixed gear but does apply to amphibious rotorcraft with retractable gear.
(5) A "landing gear down" lock is required. An optional uplock may be used if it meets reliability requirements.
(6) A (ground) operation test should be conducted to ensure proper functioning of the system.
(7) Location and operation of the control lever or device must comply with § 27.777. This section includes identification of controls to prevent confusion and inadvertent operation. Amendment 27-21 added new § 27.779 for motion and effect of cockpit controls. Specifically, § 27.779(c) pertains to motion and effect of normal landing gear controls. Section 25.781 of Part 25 contains large airplane design requirements for motion, effect, and shape of cockpit controls and their knobs and should be consulted for further guidance.
(8) A landing gear warning is required as prescribed in Paragraph (g). Certain features are required. The landing gear shall be extended and locked.
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 § 27.729(a)(1), (a)(2), and (a)(3).
(2) Wheel well doors, if installed, should be designed for the aerodynamic loads, including loads from yawing conditions (angles selected by the applicant) 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.
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