The critical condition for VNE controllability testing is ordinarily aft c.g., MC power, and minimum power-on rotor rpm, although power and rpm variations should be specifically evaluated to verify their effects.
turbine engine is sensitive to ambient temperatures which affect the engine's ability to produce rated maximun continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit maximum continuous power would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward c.g. should be spot checked for any "tuck under" tendency at high speed. The VNE controllability test is normally accomplished shortly after the 1.1 VNE (or 1.1 Vn) point obtained during stability tests required by § 29.175(b). Controllability must be satisfactory for both conditions. If VNE varies with altitude or temperature, VNE for existing ambient conditions is utilized for the test. Extremes of the altitude/temperature envelope should be analyzed and investigated by flight test.
(3) The critical condition for controllability testing in a hover is ordinarily forward c.g. at maximum weight with minimum power-on rotor rpm. For rearward flight testing of configurations where the forward c.g. limit varies with weight, low or high gross weight may be critical. Lateral c.g. limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds.
Testing in calm winds is preferred. Hover controllability testing should be accomplished with the lowest portion of the helicopter at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the helicopter is held in the most critical position relative to wind.
(4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of his engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and they should be specified in the TIA. Particular attention should be given to minimum stablized and minimum transient rotor rpm limits. These values must be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On Category A installations the maximum power output of each engine must be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and over-temping. This is needed for compliance with § 29.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation; however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers. Substantiation of this feature should be accomplished primarily by engine and drive system ground tests.
Chap 2 Par 81 155
AC 29-2, CHG 1 11/19/84 * (5) Longitudinal cyclic authority at VNE with any power setting must permit suitable nose down pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft's response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral oyolic to correct roll in conjuction with forward cyclic to correct pitchup is satisfactory. Obviously during the conduct of these tests, all available However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion. techniques should be utilized when the pilot finds himself "out of control."
(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all engine inoperative requirements of § 29.75(b) (5) and (c). See paragraph 67 of this advisory circular.
82. § 29.161 (through Amendment 29-19) TRIM CONTROL.
a. Explanation.
The pilot has many tasks to perform with each hand during sustained flight conditions. The trim requirement is intended to provide the pilot with a reference cyclic control position for the given flight condition, reduce the physical demands to maintain a given flight condition, and allow the pilot to release the cyclic control for brief periods of time to perform other cockpit duties. A primary flight control which can move when released imposes an additional pilot workload by requiring a continuous hands-on condition. It is not intended to require that control forces be reduced to zero by the trim control during dynamic maneuvers such as takeoff acceleration.
characteristics. One popular method of meeting this requirement is through the use of control balance springs in conjunction with a small amount of built-in control system friction. Other methods include use of friction, magnetic brakes,
(3) This regulation is not intended to require zero friction or zero breakout force in the control system, nor is it intended to require automatic control'recentering. The regulation, in fact, specifically prohibits excessive high friction or high breakout forces which would produce undesirable discontinuities in the primary control force gradient.
Chap 2 156 Par 81
* 140. § 29.351 (through Amendment 29-19) YAWING CONDITIONS.
a. Explanation. The rule requires proof of a rotorcraft "structural" yaw 500331 or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to VH or VNE, whichever 18 less.
structurally safe for the thrust capability of the directional control system.
(1) The rotorcraft structure must be designed to withstand the loads for the specified yaw conditions. The standard does not require a structural flight demonstration. It is a structural design standard.
(2) Maximum displacement of the directional control, except as limited by pilot effort (130 pounds; § 29.397(a)), is required for the conditions cited in the rule. A control system rate limiter or a yaw damper may be used. The total displacement is therefore a function of time as well as the maximum effort applied (130 pounds).
(1) At low airspeeds, 90° yaw (sideward flight) should be the design limit.
(11) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.
(111) At high airspeeds, the maximum tail rotor thrust will be combined with the vertical (directional) stabilizer surface load, if a stabilizer is used, as specified by § 29.351 (b) (1).
(iv) At high airspeeds, while the rotorcraft is in the sideslip condition, the directional control is then returned to the neutral position, attendent with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stablizer.
(v) Both right and left yaw conditions should be proven.
(3) The tail rotor attachment structure must comply with § 29.403.
(4) The vertical stabilizing surface must also comply with § 29.413.
b. Procedures•
(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Model FH1100. The Hiller FHl100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.
(1) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.
Chap 2 Par 140 235
AC 29-2, CHG 1 11/19/84 * (11) For vertical stablizers that house the tail rotor output gearbox, such as the Hiller Model FH1100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration NOTE: For one example: At stablized yaw to the right (left pedal depressed to limit) (§ 29.351(b) (2)), the tail rotor thrust moment should equal the restoring moment of the tail boom, vertical stabilizer and main rotor torque. As stated by § 29.351(b)(3), the tail rotor thrust moment then i9 added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 29.351(b) (3)) and vertical stabilizer load is generally one of the critical design conditions for the fuselage/tail boom.
(i1i) For vertical stablizers or fins that have an offset incidence angie with respect to the helicopter axis, the vertical fin moment is added, or subtracted as applicable, to the tail rotor thrust moment. The condition stated in § 29.351(b)(1) may result in adding the fin load to the tail rotor thrust.
(Iv) Low airspeed maneuvers, such as sideward, rearward, and hover turns over a spot, typically impose insignificant aerodynamic loads on the fugelage and/or tail boom. The aerodynamic loads at VH or VNE, whichever is required, are generally the significant aerodynamic design loads.
A rational assessment of the various yaw conditions may be used to reduce the load deviation and analysis to the critical rotorcraft design conditions.
(vi) The helicopter structure shall be analyzed or tested for loads derived from the critical design conditions.
(vi1) A simple structural design envelope may be derived from these design data. If the right or left yaw limits are not very different, common, conservative design limits may be used. A sample yaw/forward speed diagram, as derived from design analysis of the characteristics of a hypothetical helicopter, is presented in Figure 140-l. A table of values would also suffice. This figure reflects characteristics which include a 90° yaw when the directional control inputs are applied at low airspeeds (up to 30 knots presumably the maximum are applied at VH, with a straight line variation from 30 knot forward apeed to • sideward flight speed of which this aircraft is capable) and 10° yaw when they Chap 2 236 Par 140
90° -90° Yaw Degrees 10° Max.
30 knot sideward Flight Airspeed FIGURE 140-1 * (vii1) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both a visual readout for the pilot and a record, such as an oscillograph trace, have been used. This test may be conducted in the flight test program or in the flight load survey program. This record should confirm the yaw angle used in design as conservative with respeot to operational and actual flight characteristics. This test is not a requirement however.
Chap 2 Par 140 237
AC 29-2, CHG 1 11/19/84
141. § 29.361 (through Amendment 29-19) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft should be designed for limit engine to que values, as prescribed by the rule, to account for maximum engine torque, including transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.
(2) A factor of 1.25 applies to maximum continuous power for turbine engines.
(3) Torque factors are also specified for reciprocating engines having two or more cylinders § 29.361a.2. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.
(4) Paragraph 206 of this document concerns $ 29.549(e) that contains standards associated with 2 1/2-minute power rating.
D• Procedures•
(1) The engine torque associated with the maximum continuous power condition must be multiplied by the appropriate torque factor to obtain the engine torque value used for structural substantiation purposes of the rotorcraft.
(2) The torque values associated with the minimum power-on rpm limit should be used. Maximum power-on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic.
142.-151. RESERVED.
Chap 2 238 (thru 248) Par 141
SECTION 7. STRENGTH REQUIREMENTS - GENERAL * 121. § 29.301 (through Amendment 29-19) LOADS.
a. Explanation,
(1) The rule is a general statement concerning limit and ultimate loads, and the application of these loads to the rotorcraft.
(2) Ultimate loads are limit loads multiplied by the prescribed factors of safety.
(3) The specified loads must be distributed appropriately or conservatively and signficant changes in distribution of the loads, as a result of deflection, must be taken into account.
b. Procedures. The design criteria report and/or design loads report must contain data that comply with the rule.
122. § 29.303 (through Amendment 29-19) FACTOR OF SAFETY.
a• Explanation.
(1) Unless otherwise provided by FAR Part 29, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will assure that the design strength of the rotorcraft is greater than the design loads contained in FAR Part 29.
(2) Other rules, $$ 29.561 (b)(3) and 29.787(c), specify use of defined ultimate inertia forces for protection of occupants.
b. Procedures.
(1) The design criteria report and/or design loads report must contain data that include the appropriate factor of safety.
(2) The factor of safety multiplies the limit external and inertia loads. The rule does allow the application of this factor to the resulting 'limit internal" stresses if it is more conservative.
123. § 29.305 (through Amendment 29-19) STRENGTH AND DEFORMATION.
a. Explanation.
(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.
(2) The structure must support, in a static test, ultimate loads for 3 seconds without failure or dynamic tests simulating actual load application may * be used.
Chap 2 Par 121 197
AC 29-2, CHG 1 11/19/84
(3) Section 29.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by struotural analysis.
See paragraph 124 of this advisory circular.
b. Procedures. Any test results, static or dynamic, must satisfy the limitations or acceptance criteria contained in the rule.
(1) Any test proposals submitted for approval that are used to demonstrate compliance with sections of FAR Part 29 must contain the criteria stated in the rule.
(11) Any test results reports shall contain data and information showing the test results comply with the standard.
Chap 2 198 Par 123
124. § 29.307 (through Amendment 29-19) PROOF OF STRUCTURE.
a• Explanation.
(1) The rule requires compliance for each critical loading condition.
Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in subparagraph (b) (6).
(2) Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable.
(3) Fatigue substantiation requirements are explained further in paragraph 230, § 29.571 of this advisory ciroular.
b. Procedures.
(1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the oritical loading conditions will be selected for analytical substantiation in struotural (static and fatigue) reports and dynamics (vibration and stability) reports and fatigue, statio, dynamio, or operational test reports.
(2) Whenever tests are used or required, a test proposal or plan must be approved prior to the tests. The test article must have received conformity inspections and must have been acoepted by the FAA for the test. Test fixtures and instrumentation must also be acceptable to the FAA (using DER's as appropriate) prior to start of the test. The quality control office of the applicant or other qualified pesonnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA or DER performing this task. The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings, such as test fixture details, or layering of approvals is not intended or envisaged by this policy. Drawings, sketches or photographs have been used by the FAA to control and to assure correct location, direction, and magnitude of loads and other oritical test parameters.
(3) Structural analysis has been accepted for rotorcraft in place of static tests. Generally the helicopter airframe should have frequency placements remote to predominate rotor excitation sources, including rotor harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under § 29.571, critical loaded areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data may be compared to the analytical data to verify accuracy.
Chap 2 Par 124 199
AC 29-2, CHG 1 11/19/84
(4) Subparagraph (b) of the rule specifies certain tests. Test proposals must be approved prior to conducting official FAA tests. Other paragraphs in this advisory circular pertain to those tests.
125. § 29.309 (through Amendment 29-19) DESIGN LIMITATIONS.
a. Explanation.
(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant must establish these structural limitations to facilitate design of the rotorcraft.
(2) Refer to the rule for the specifio requirements.
b. Procedures.
(1) The design criteria and/or design load report should contain the design limits specified.
(2) These items are structural design limits. Other requirements may result in narrowing the ranges of type design limits or in reducing limits. It is not necessary to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits.
126.-135. RESERVED.
Chap 2 200 (thru 230) Par 124
AC 43.13-1A CHG 1 Page 165 1975 Chapter 10. HYDRAULIC AND PNEUMATIC SYSTEMS Section 1. HYDRAULIC SYSTEMS
392. GENERAL. Maintain, service, and adjust connection fittings: AN-818 nut and AN-819 airplane hydraulic systems in accordance with sleeve.
manufacturers' maintenance manuals and pertinent component maintenance manuals. Certain Minimum bend radil * general principles of maintenance and repair for tightening tube nuts Wrench torque range measured at Inside of bend.
which apply are outlined below. Tube O.D. (Inch pounds) Dimenaton in inches. * (inches)
393. HYDRAULIC LINES AND FITTINGS. Carefully Alum. alloy 1100-H14, Alum. alloy 1100-H114, inspect all lines and fittings at regular intervals 5052-0 Steel 5052-0 Steel to insure airworthiness. Investigate any evidence of fluid loss or leaks. Check metal lines for 30-70 leaks, loose anchorages, scratches, kinks, or other ¾а 40-65 60-80 50-90 70-120 21/32 damage. Inspect fittings and connections for 75-125 00-150 1⅜ leakage, looseness, cracks, burrs, or other damage. 150-250 155-250 15¾8 1¾в Replace or repair defective elements. 200-350 300-400 1¼ 1¾ 300-500 430-575 1⅓ 218
a. Replacement of Metal Lines. When inspec- 500-700 550-750 1¾ 2% tion shows a line to be damaged or defective, 1 600-900 3 replace the entire line or if the damaged section 1¼ 600-900 9% is localized, a repair section may be inserted. 14½ 5¼ In replacing lines, always use tubing of the same 1¾ 2 size and material as the original line. Use the old tubing as a template in bending the new line, FIGURE 10.1.-Tube data.
a template can be made from soft iron wire.
Soft aluminum tubing (1100, 3003, or 5052) under ¼-inch outside diameter may be bent by hand. For all other tubing use an acceptable hand or power tube bending tool. Bend tubing carefully to avoid excessive flattening, kinking, or wrinkling. Minimum bend radii values are shown in figure 10.1. A small amount of flattening in bends is acceptable but do not exceed an amount such that the small diameter of the flattened portion is less than 75 percent of the original outside diameter. When installing the replacement tubing, line it up correctly with the mating part so that it is not forced into line by means of the coupling nuts.
b. Tube Connections. Many tubing connections are made using flured tube ends, and standard In forming fares, cut the tube ends square, file smooth, remove all burrs and sharp edges, and thoroughly clean. The tubing is then flared using the correct 37° aviation flare forming tool for the size of tubing and type of fitting. A double flare is used on soft aluminum tubing ⅜-inch outside diameter and under, and a single flare on all other tubing. In making the connections, use hydraulic fluid as a lubricant and then tighten. Overtightening will damage the tube or fitting, which may cause a failure; undertightening may cause leakage which could result in a system failure.
Mistaken use of 45° automotive flare forming tools will result in improper tubing flare shape and angle causing misfit, stress and strain, and probable system failure.
Chop 10 Par 392
c. Repair of Metal Tube Lines. Minor dents and scratches in tubing may be repaired. Scratches or nicks no deeper than 10 percent of the wall thickness in aluminum alloy tubing, that are not in the heel of a bend, may be repaired by burnishing with hand tools. Replace lines with severe die marks, seams, or splits in the tube.
Any crack or deformity in a flare is also unacceptable and cause for rejection. A dent less than 20 percent of the tube diameter is not objectionable unless it is in the heel of a bend.
Dents may be removed by drawing a bullet of proper size through the tube by means of a length of cable. A severely damaged line should be replaced; however, it may be repaired by cutting out the damaged section and inserting a tube section of the same size and material. Flare both ends of the undamaged and replacement tube sections and make the connection by using standard unions, sleeves, and tube nuts. If the damaged portion is short enough, omit the insert tube and repair by using one union and two sets of connection fittings.
d. Replacement of Flexible Lines. When replacement of a flexible line is necessary, use the same type, size, and length of hose as the. line to be replaced. If the replacement of a hose with swaged-end-type fittings is necessary, obtain a new hose assembly of the correct size and composition. Certain synthetic oils require a specially compounded synthetic rubber hose which is compatible. Refer to the aircraft manufacturer's service information for correct part number for replacement hose. If the fittings on each and are of the collet type or sleeve type in replacement may be fabricated as shown figure 10.2. Typical aircraft hose specifications and their uses are shown in figure 10.3. Install hose assemblies without twisting. (See figure
10.4.) Never stretch a hose tight between two fittings as this will result in overstressing and eventual failure. The length of hose should be sufficient to provide about 5 percent to 8 percent slack. Avoid tight bends in flex lines as they may result in failure. Never exceed the minimum bend radii as indicated in figure 10.5.
Teflon hose is used in many aircraft systems because of its superior qualities for certain applications. Teflon is compounded from tetrafluoroethylene resin which is unaffected by fluids Par 393 AC 43.13-1A CHG 1 1975 normally used in aircraft. It has an operating range of -65° F. to 450° F. For these reasons, Teflon is used in hydraulic and engine lubricating systems where temperatures and pressures preclude the use of rubber hose. Although Teflon hose has excellent performance qualities, it also has peculiar characteristics that require extra care in handling. It tends to assume a permanent set when exposed to high pressure or temperature. Do not attempt to straighten a hose that has been in service. Any excessive bending or twisting will cause kinking or weakening of the tubing wall. Replace any hose that shows signs of leakage, abrasion, or kinking.
Any hose suspected of kinking may be checked with a steel ball of proper size. Figure 10,6 shows hose and ball sizes. The ball will not pass through if the hose is distorted beyond limits.
If the hose fittings are of the reusable type, a replacement hose may be fabricated as described in figure 10.2. When a hose assembly is removed the ends should be tied as shown in figure 10.7, so that its preformed shape will be maintained. Refer to figure 10.9 for minimum bend radii of Teflon hose.
All flexible hose installations should be supported at least every 24 inches. Closer supports are preferred. They should be carefully routed and securely clamped to avoid abrasion, kinking, or excessive flexing. Excessive flexing may cause weakening of the hose or loosening at the fittings.
e. O-Ring Seals. A thorough understanding of O-ring seal applications is necessary to determine when replacement must be made. The simplest application is where the O-ring merely serves as a gasket when it is compressed within a recessed area by applying pressure with a packing nut or screw cap.
Leakage is not normally acceptable in this type of installation. In other installations the O- ring seals depend primarily upon their resiliency to accomplish their sealing uction. When moving parts are involved, minor seepage may be normal and acceptable. A moist surface found on moving parts of hydraulic units is an indication the seal is being properly lubricated.
In pneumatic systems, seal lubrication is provided by the installation of a grease-impreg- Chap 10