The test loads will be applied and reacted using mating aircraft parts or special fixtures which simulate the mating parts. The methods and apparatus to be used for the static tests of the castings are discussed in the apparatus and method section of this report.
Testing will be conducted in... (location).
TEST SPECIMEN The castings which will be tested are listed in numerical order in figure AC 27.621-2.
Those castings which, after structural analysis, show less than a 1.5 casting factor will be tested. All directions are given with reference to a forward facing position in the rotorcraft.
On the basis of a radiographic examination, the three castings which are of the poorest acceptable quality in the first production lot of castings will be selected as test specimens. The poorest of the three castings will be selected as the initial test casting and its radiograph or ASTM standard will be used as the standard for accepting future castings of the particular part unless later standards are approved. Three castings must be tested for each critical condition for each part.
Conformity Inspection Each machined casting will be subjected to an FAA/AUTHORITY conformity inspection prior to testing to determine compliance with the type design drawings. A conformity report for each casting may be incorporated in Part II, Test Results, of this report.
The test specimen will be permanently marked or defaced after testing to preclude its use on a rotorcraft.
See figure AC 27.621-2 for an example of a convenient means of listing castings.
TEST LOAD The test load(s) to be applied to each casting represents the critical loading condition(s) for that casting. The critical conditions on each of the castings were determined by the design criteria and substantiating data approved by the FAA/AUTHORITY.
The design criteria for all of the castings to be static tested may fall into one of two categories. The load factors and structural acceptability requirements for each category are discussed below. Casting factors that are included on the load sheets of each part do not apply in the discussion below. (See paragraph AC 27.621b(2) for casting Page D - 21
Castings Designed to Limit Load Conditions A structural analysis of each test casting showing the critical design limit load conditions is given in the data (reference report number here). The load factors for the static test of the castings are as follows:
1.15 x design limit load = design yield load
1.50 x design limit load = design ultimate load Castings Designed Only to Crash Landing Conditions The castings in this category were designed using a crash landing load factor for the design ultimate load. The design yield load criteria of 1.15 x limit load need not apply to these castings. The test loads for these castings may be given in terms of design ultimate load on the individual casting load sheets shown in Part I of this report.
Test Procedures Depending on the results of the initial static test of each casting, the following procedure will be used.
a. If in the initial test of critical castings the casting is found to have a casting factor of 1.5 (1.5 x design ultimate load), the casting will be considered acceptable and no further tests will be conducted.
b. If in the initial tests) the critical casting is found to have a casting factor less than 1.5 but equal to or greater than 1.25, two additional castings will be tested for each critical load condition. Each must also show a minimum casting factor of 1.25.
c. If in the initial test, or in one of two additional tests, a casting shows a casting factor less than 1.25 times design ultimate or yields prior to reaching 1.15 times design limit load, the casting will be redesigned and retested. The yield criteria are also applicable to the first two procedures with the exception of critical castings designed to crash landing conditions.
TEST APPARATUS AND METHOD The Model XYZ casting static tests will be conducted using fixtures designed to simulate the installation of the castings in the aircraft. Where practical, mating aircraft parts will be used to apply and react test loads. When practical, the static tests will be conducted with mating castings assembled when the critical loads for the mating castings are compatible; otherwise, fixtures simulating the mating parts may be designed and fabricated for the tests. Assembly hardware used to mount test castings will be the same as hardware used on the rotorcraft. All bolt torques and other assembly notes will conform to the type design assembly instructions.
Page D - 22
The tests will be conducted using calibrated load measuring devices such as hydraulic cylinders and pressure gages, load cells, strain gage bridges, or dead weights.
Deflections of the casting may be measured using graduated dial indicators or scales in all tests. The deflection indicators will be based or mounted on the casting and will measure casting deflection only when possible; otherwise, the indicators will be based on the fixture and measure deflection of the casting relative to the fixture. Deflection readings will be made at 20 percent increments of limit load through 100 percent of limit load and at 115 percent of limit load. These increments may be changed if necessary.
Permanent deformation readings will be made after relieving 115 percent and 150 percent of limit load.
See figure AC 27.621-1 as an example of a load sheet.
Page D - 23
ULTIMATE LOADS -1,000 # COMP *5,000 # TENS Tension Comp.
Tension Reaction Compression Reaction FIGURE AC 27,621-1
EXAMPLE OF CASTING LOAD SHEET
RETRACT ACTUATOR SUPPORT - LANDING GE Page D - 24
Include spherical bearing with clamped-up bolt and a link in the test setup to confirm the stability. Loads are based on a jam condition with actuator operating at 1,700 PSI pressure maximum.
A 1.25 casting factor is included in these loads.
These loads were derived from data in approved structural loads and analysis report.
END OF SAMPLE REPORT
(4) The format of the previous guidance material may be changed to accommodate the applicant's method of data presentation.
(5) Nonstructural castings may be tested and included in the test report.
(6) Cast fluid containers, including hydraulic fluid containers, may be tested as prescribed in other rules of FAR Part 27 and a test proposal and test results report may be included in the casting test report, or an appropriate report may be referenced for convenience. We recommend use of one report to contain test data or reference to test data for all castings used on the rotorcraft.
Page D - 25
AC 27 1B 9/30/99 FIGURE AC 27.621-2 EXAMPLE
CASTINGS TO BE STATIC TESTED FOR MODEL XYZ REF. LOAD
MACHINE OR SHEET CASTING NO. ASSY. NO. NAME AND LOCATION MATERIAL FIG. NO.
Base Assembly, Pilot's Collective Column Page D - 26
AC 27.623. $ 27.623 BEARING FACTORS.
a. Explanation.
(1) The rule requires use of a minimum bearing factor in free fit joints to account for effects of typical relative motion. A minimum value is not specified in the rule. The factor, appropriate for the application, is applied to the ultimate bearing strength of the softest material used as a bearing. A definition of free fit (clearance fit) is noted in paragraph AC 27.623b(7) below.
(2) A bearing factor, appropriate for the application, shall be used unless a larger factor is used.
(3) For reference, specific bearing factors are contained in § 29.685(e) for transport rotorcraft control system joints subject to angular rotation. These factors are applied to the ultimate bearing strength of the softest material used as a bearing in the control system. Control systems ball, roller, or needle bearings are covered by § 29.685(f) for transport rotorcraft.
(4) MIL-HDBK-5D, paragraph 8.3, refers to design standards for plan or journal bearings or bushings. These standards are found in Air Force Systems Command Design Handbook AFSC DH-2-1, Airframe, Chapters 3 and 6.
b. Procedures.
(1) Control system joint bearings are discussed under § 27.685 and paragraph AC 27.685A, but the bearing factors are noted here for convenience. For transport rotorcraft control systems, § 29.685(e) requires a 2.0 bearing factor for cable systems and a 3.33 bearing factor for push-pull systems other than ball and roller bearing systems. The manufacturer's static, non-Brinell rating of ball and roller bearings should not be exceeded. Use of this for normal category rotorcraft is recommended.
(2) A landing gear pivot, grease-lubricated, plain bearing is one example of a free fit subject to pounding or vibration. A bearing factor of 2.0 may be used or another factor may be proven for a grease-lubricated plain bearing or bushing to account for the anticipated higher loads caused by pounding or vibration. See paragraph AC 27.623b(6) for recommendations on ball or roller bearings.
(3) A typical engine mount bolt installation with a plain bearing having a free or loose fit (not interference fit) is another example of a sleeve bearing application subject to a design bearing factor. As an EXAMPLE OR ILLUSTRATION, a bearing factor of
1.85 may be applied to the design loads on the softest material reacting the bearing loads. A different but appropriate factor will be acceptable. The design limit load may be calculated for the example of a 0.312-inch-diameter bolt in a 2-inch-long bearing.
The bearing projected area is 0.312 × 2 = 0.624-square-inches. The design limit load is 3,000 pounds. The design limit bearing stress is 3,000 pounds/0.624-square-inch x Page D - 27
1.85 = 8,894 PSI. If a free or loose fit is not used; i.e., tighter than free fit, a bearing factor is not required. See paragraph AC 27.623a(4) for bearing factors.
(4) Military standard part specifications, MS 21240, "Bearing, Sleeve Plain, TFE Lined," and MS 21241, "Flanged Bearing, Sleeve Plain, TFE Lined," contain allowable load ratings, static and dynamic, that apply to the particular use of the bearing. An appropriate bearing factor should be applied to the static rating. Military Specification (MIL-B-8943A, Amendment 3, "Bearing, Sleeve, Plain, and Flanged, TFE Lined" (temperature range -65° F to +250° F) shows that MS 21240 and MS 21241 sleeve bearings have been superseded by MS 1934/1 and MS 81934/2 sleeve bearings, respectively. Military Specification MIL-B-81934, Amendment 2, "Bearings, Sleeve, Plain and Flanged, Self-Lubricating," uses TFE liners. These bearings are intended for use in a temperature range from -65° F to +325° F. Whenever a sleeve bearing is used, an appropriate bearing factor should be applied to the static rating that is contained in the specification or standard. Other sleeve bearings are contained in standards NAS 72 through NAS 77, NAS 537, and NAS 538. The installation design information is only contained in standards NAS 72 through NAS 74. These types of plain sleeve bearings are designed for clamping to the shaft or bolt with relative motion occurring on the bearing outside diameter. An appropriate bearing factor is required for the application.
(5) The minimum fitting factor 1.15, specified by § 27.625, must be applied as specified to account for load distribution at the fitting. This fitting factor need not apply to plain or journal "bearings" whose "bearing factor" exceeds 1.15.
(6) For airframe and landing gear structural joints, the manufacturer's static, non-Brinell rating of ball and roller bearings should not be exceeded. ABEC Class 1 bearings or better quality bearings may be used in airframe structural joints and landing gear; ABEC Class 3, 5, or 7 bearings should be used in rotor pivot joints. The non-Brinell rating includes consideration of the bearing factor, and no other bearing factor is necessary.
(7) A free fit was described in American Standards Association (ASA) Standard B4a-1925. The "free fit" clearances and tolerances of this old standard are now called Class RC6, Medium Running Fit, in ASA Standard B4.1, 1955. As an illustration using these standards, a 1-inch diameter shaft and a plain sleeve bearing would have a clearance ranging from 0.0014 to 0.0040 inch.
AC 27.625. § 27.625 FITTING FACTORS.
a. Explanation. A 1.15 factor is specified to ensure that the calculated load and stress distribution within any fitting is conservative. Application of the factor is excluded or is an exception as stated in the rule.
b. Procedures.
(1) The factor may be applied to the calculated load or stress for the fitting.
Page D - 28
(2) The structural design substantiating data should include the fitting factor and where applicable should include, but not be limited to, the rotor system. The rotor system includes the rotor blade attachments, rotor head and hubs, and boosted control system elements. Other typical areas that may be considered are tail rotor gearbox attachment, tailboom to fuselage fittings, transmission pylon attachments, and landing gear attachment to the rotorcraft.
(3) The fitting factor is not required in the following applications:
(i) Joints such as continuous joints in metal plating, welded joints and scarf joints in wood.
(II) Elements proven by limit or ultimate load tests such as nonboosted control system parts.
(iii) Elements for which a larger load factor is used such as a casting factor, a 1.33 retention factor when required for seats and safety belts, a fatigue factor, bearing factor or special factor greater than 1.15, crash load factors that are the only design case, and crash load factors that exceed limit load factors x 1.5 x 1.15.
(iv) Elements for which the failure mode does not affect safety of flight or occupant safety.
AC 27.625A § 27.625 (Amendment 27-35) FITTING FACTORS.
a. Explanation. Amendment 27-35 added § 27.625(d) that requires a 1.33 factor applied to the emergency landing loads of § 27.561(b)(3) for the substantiation of attachments of each seat, berth and litter to the structure and each safety belt or harness to the seat, litter, or structure.
b. Procedures. All of the advisory material pertaining to this section remains in effect with the following additions.
(1) A fitting factor of 1.33 must be applied to the emergency landing loads of § 27.561(b)(3) when evaluating the attachments of the seat, berth, and litter to the structure, and each safety belt and harness attachment to the seat, berth, litter, or structure.
(2) The 1.33 factor is required whether analysis or test is used.
AC 27.629. $27.629 FLUTTER.
a. Explanation. The rule requires that the rotorcraft "be free from flutter under each appropriate speed and power condition."
Page D - 29
special aroidure t. Freades, frs, fnd sais ta be shown for the entire rotorcraft with
(1) Flutter is defined as an aeroelastic instability resulting primarily from coupling of flap and pitch bending modes.
(2) Freedom from flutter may be shown by analysis or by appropriately instrumented flight flutter tests.
(3) The flight load survey proposal submitted for compliance with § 27.571 may also contain tests to fulfill compliance with § 27.629.
(4) Flight loads survey data or flight flutter test data should be reviewed to ensure that excessive oscillatory deflections of rotors or surfaces will not be encountered.
(5) Sensitivity analyses should be conducted to ensure that normal wear in the pitch change mechanisms of the main rotor blades and tail rotor blades does not reduce the effective stiffnesses sufficiently to cause flutter.
Page D - 30
SUBPART D - DESIGN AND CONSTRUCTION ROTORS AC 27.653 § 27.653 (Amendment 27-2) PRESSURE VENTING AND DRAINAGE OF ROTOR BLADES.
a. Explanation. The rule requires each rotor blade to be provided with venting and drainage means (i.e., holes, etc.) or else the blade must be sealed and designed to withstand internal pressure.
b. Procedures. Although the rule provides for venting and drainage features, recently certificated blades have been designed to be sealed and to sustain the "maximum pressure differentials expected in service." For modern blade designs, the internal pressure buildup due to environmental effects and centrifugal acceleration effects (near the tip) can be readily sustained with moisture sealing accomplished. The use of sealed blades is highly advantageous and recommended because of the possibility for severe corrosion damage resulting from trapped moisture and because of the difficulty in finding internal corrosion damage by use of field level inspections.
AC 27.659 § 27.659 (Amendment 27-2) MASS BALANCE.
a. Explanation. The rule requires that mass balancing of rotors and blades be provided, as necessary, to prevent excessive vibration and flutter. Further, the rule requires structural substantiation of the mass balance installation.
b. Procedures.
(1) The weight, geometry, and location of rotor and blade mass balance devices are determined as the requirements of §§ 27.571 and 27.629 are met.
(2) The structural substantiation should show static strength to meet the maneuver and gust loads of §§ 27.337, 27.339, and 27.341. In addition, the main rotor loads of § 27.547(c) should be substantiated. The fatigue strength of the mass balance devices (including structural supports) should meet the requirements of § 27.571.
(3) In addition to the appropriate strength requirements, some recent designs have included features which trap the balance weight inside a limited area even if the primary attachment means (adhesive, bolts, etc.) fail. This type of design feature is recommended because of the severe loading environment to which balance devices are subjected.
AC 27.661 § 27.661 (through Amendment 29-2) ROTOR BLADE CLEARANCE.
a. Explanation.
Page D - 31
(1) This paragraph discusses the regulatory requirement contained in § 27.661.
That requirement is that there must be enough clearance between the rotor blades (main and tail rotor blades) and other parts of the structure to prevent the blades from striking any part of the structure during any operating condition.
(2) In the past, some rotorcraft that have been shown to comply with § 27.661 during the certification process have experienced subsequent accidents involving inflight contact between the main rotor and airframe (rotor/airframe contact). Completion of developmental and TIA flight testing without a rotor/airframe contact incident has proven not to be adequate demonstration of compliance with § 27.661 in all cases.
(3) Historically, in-flight rotor/airtrame contact accidents have occurred as a result of mast bumping, rotor stall, or excessive rotor flapping due to control manipulation. For some rotorcraft, a more thorough examination may be required to ensure adequate clearances.
b. Procedures. Testing should be conducted by the applicant, prior to FAA/AUTHORITY participation, to ensure that the rotorcraft is in compliance with § 27.661 in all areas of the envelope during all operational maneuvers expected throughout the life of the aircraft. The tests should be performed concurrently with performance, flight characteristics, and flight loads testing. Tests should include:
(1) A blade flapping survey to determine flapping angles/margins, blade bending, and blade clearance from the entire airframe. Data may be gathered from instrumented flapping hinges, instrumented blades, high-speed video from airframe mounted cameras, a chase aircraft, or other acceptable means.
(2) Determine that margin exists between the minimum rotor RPM encountered during testing for compliance with § 27.143(d) and the RPM (power off) at which analysis shows that the rotor will experience a significant stall. A significant stall condition may be defined by the rotor reaching an RPM from which normal operating RPM is unrecoverable due to drag on the main rotor blades or, a stall that results in excessive main rotor flapping. The rotor RPM decay rate under the critical conditions of weight, density altitude, minimum approved power-on rotor RPM must provide a margin between the minimum rotor speed achieved during demonstration of compliance with §§ 27.79 and 27.143(d) and the analytically derived rotor stall RPM for the same conditions. For example, the minimum rotor RPM resulting during H-V tests must allow for a margin above the rotor stall value to allow for variations that may occur during operational flying.
(3) During parts of the certification flight test program, frangible devices (wood dowels) or other means of measuring clearance, may be requested to confirm that the clearances shown in the drawings and verified during company flight tests are adequate in all operating conditions. Balsa wood dowels or styrofoam pads may be clamped to the aft part of the fuselage and cabin roof within the rotor arc. Such devices may be Page D - 32