(2) Gust loads for any vertical stabilizing surface should be derived for lateral or sideward gusts, as well as the head-on horizontal gusts. See Paragraph No. 159, § 29.413(a)(2) of this advisory circular.
(3) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward, as well as for head-on gusts. See Paragraph 159.
b. Procedures.
(1) Either sharp-edged (instantaneous) gusts or sharp-edged gusts modified by an alleviation (attenuation) factor may be used for calculating aerodynamic loads for the rotorcraft and any installed stabilizing surfaces. The following conditions may be used:
Vertical gusts may be considered normal to the flight path of the rotorcraft except during hover or low speed flight (20 knots or less) when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.
(il) For a vertical stabilizing surface, the horizontal gusts are normal to the flight path of the rotorcraft except during hover or low speed flight when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.
(III) A primary effect of encountering the gust is to change the lift of the rotors and rotorcraft surfaces. Of primary concern is the gust load or lift created by the main rotor or rotors. The lift increment of the horizontal stabilizing surface and fuselage are generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.
(iv) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.
(v) The gust velocity may be assumed uniform across the rotorcraft.
(vi) Gust loads on the stabilizing surfaces are required as stated in Paragraph 159 of this advisory circular.
(2) The rotorcraft design maneuvering load factors may generally exceed the design gust load factors calculated in compliance with this rule. This may be attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to assure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load tactor.
(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA - Ames Specialist's Meeting on Rotorcraft Par 139 243
Dynamics, February 13-15, 1974. The paper, entitled, "Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects," was written by P.J.
Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.
140. § 29.351 YAWING CONDITIONS.
a. Explanation. The rule requires proof of a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to Vu or VNE, whichever is less. The rotorcraft must be 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).
(i) At low airspeeds, 90° yaw (sideward flight) should be the design limit.
(ii) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.
(ill) 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, attendant with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stabilizer.
(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 244 Par 139
as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tall rotor output gearbox.
For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.
(її) For vertical stabilizers 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 stabilized 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 is 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.
(iii) For vertical stabilizers or fins that have an offset incidence angle with respect to the rotorcraft 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 fuselage and/or tail boom. The aerodynamic loads at VH or VNE, whichever is required, are generally the significant aerodynamic design loads.
(v) 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 rotorcraft structure shall be analyzed or tested for loads derived from the critical design conditions.
(vii) 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 rotorcraft, is presented in Figure 140-1. 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 sideward flight speed of which this aircraft is capable) and 10° yaw when they are applied at VH, with a straight line variation from 30 knot forward speed to VH.
Par 140 245
90° 90° YAW DEGREES 10° MAX.
0° AIRSPEED 30 KNOT LESSER OF
SIDEWARD VH OR VNE FLIGHT FIGURE 140-1 SAMPLE YAW/FORWARD SPEED DIAGRAM 246 Par 140
(vili) 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 respect to operational and actual flight characteristics. This test is not a requirement however.
140A. § 29.351 (Amendment 29-30) YAWING CONDITIONS.
a. Explanation. Amendment 29-30 adds maximum sideslip angles to the existing § 29.351 for structural design purposes. The standard should apply to power-on conditions; not power-off, since Vu is a part of the standard. For airspeeds up to 0.6 VNE, sideslip angles larger than 90° (or sideward flight) need not be considered. For airspeeds at VNE or Vu (whichever is less), sideslip angles larger than 15° need not be considered.
b. Procedures.
(1) All of the policy material pertaining to this section remains in effect with the addition of the maximum sideslip limits of 90° and 15° specified above. The rotorcraft does not need to be capable of attaining these conditions. A revised yaw/forward speed diagram is presented in Figure 140-2.
(2) FAR § 29.351(b)(1) incorrectly references § 29.395(a) for maximum pilot forces. The correct reference should be § 29.397(a).
Par 140 247
90°
YAW DEGREES (LONGITUDINAL AXIS) • 15° MAX.
0.6 VNE AIRSPEED
LESSER OF
VA OR VNE FIGURE 140-2. SAMPLE YAW/FORWARD SPEED DIAGRAM 248 Par 140
140B. § 29.351 (Amendment 29-40) YAWING CONDITIONS.
a. Explanation. Amendment 29-31 deleted FAR § 29.403 and § 29.413 since the references and requirements are adequately addressed in §§ 29.337, 29.339, and
29.341. Paragraphs 157 and 159 of this AC are retained as guidance information.
b. Procedures. All of the policy material pertaining to this section remains in effect except for references to §§ 29.403 and 29.413 in Paragraphs 140a(3) and (4) of this AC.
141. § 29.361 (Amendment 29-26) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft should be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.
A factor of 1.25 applies to maximum continuous power for turbine engines. Section 29.923 refers to torque output and § 29.927 (b) refers to other torque output conditions for use in an "endurance test."
(ii) Torque factors are also specified for reciprocating engines having two or more cylinders in § 29.361(a)(2) or § 29.361(b) of Amendment 29-26. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.
(2) Amendment 29-26 introduced additional turbine engine installation considerations for the following:
(i) Engine torque loads associated with emergency operation of governor-controlled turboshaft engines.
(ii) Torque reaction loads from sudden turbine engine stoppage which is applied to the engine and the engine suspension and restraint system.
(3) Paragraph 206 of this document concerns § 29.549(c) and (e) that contains design standards for engine mounts and adjacent structure for flight and landing and also flight with 2 ½-minute OEI power rating. Amendment 29-26 added OEI power to the standard.
(4) Section 29.547(e) (1) (il) concerns the application of limit engine torque to design of the main rotor structure.
Par 140B 249
b. Procedures.
(1) The engine torque associated with the maximum continuous power condition should 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.
(3) For turbine engines limit torque values are determined for the four cases cited. Two cases are related to "endurance" test standards.
(4) For sudden stoppage of turbine engines the engine manufacturer can reasonably provide engine rotating inertia and deceleration time expected in the event of sudden engine stoppage which generates these critical loads in the engine mounting and restraint system. These manufacturer's data should be acceptable for use in complying with this part of the design standard.
142.-151. RESERVED.
250 (thru 260) Par 141
SECTION 9. CONTROL SURFACE AND SYSTEM LOADS
152. § 29.391 GENERAL.
a. Explanation. This general rule concerns requirements for design loads of tail rotors, control or stabilizing surfaces, and their control system.
b. Procedures. The design criteria and/or the design loads report must contain the loads dictated by the referenced rules. See Paragraphs 153, 154, 155, 156, 157, 158, and 159.
152A. § 29.391 (Amendment 29-30) GENERAL.
a. Explanation. Amendment 29-30 adds an explicit reference to § 29.427, Unsymmetrical Loads (AC Paragraph 162), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 29.399, Dual Control System (AC Paragraph 155), is also added for clarification. In addition, §§ 29.401, 29.403, 29.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.
b. Procedures. The referenced AC paragraphs become 153, 153A, 154, 155, 158, and 162.
153. § 29.395 CONTROL SYSTEM.
a. Explanation. Control system design loads and the application of these loads are contained in this rule.
(1) Paragraph (a) of the rule specifies the way or means of reacting the design loads specified in §§ 29.397 and 29.399 (for dual control systems). The design loads must be imposed on any locks and stops and irreversible mechanisms in the control system. Both rotor blade horns and control surface horns must react without failure, the specified loads while the controls are in critical positions.
(2) Paragraph (b) of the rule specifies application of limit pilot forces or of the maximum loads that can be obtained in normal operation, including any single power boost system failure, whichever is greater. However, minimum limit pilot force 0.60 of the loads specified in §§ 29.397 and 29.399, may be used, as specified, in parts of the primary control system that are not stiff enough to react to the loads specified in the first part of Paragraph (b) of the rule. Note the objective for a rugged control system.
(3) Control system design feature and test requirements are found in §§ 29.671 through 29.695. Bearing factors and fitting factors are specified in §§ 29.623 and
29.625, respectively.
Par 152 261
b. Procedures.
(1) The design criteria and/or a design loads report that includes the primary control system design loads should be submitted for FAA/AUTHORITY approval.
(2) The rotorcraft control system may be tested to ultimate design loads or may be analyzed for the ultimate design loads. See Paragraph 124 of this document.
(i) It is advisable that the applicant prepare a proposal describing the procedures and techniques to be used in the static testing of the control system which reflects compliance with the condition specified. It is further advisable that the FAA/AUTHORITY concur that the tests proposed achieve that objective. Omission of these steps may result in the need for retesting. The test results should be documented.
(її) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 29.685(e), 29.623, and 29.625 must be used as specified. A structural analysis report should be used to document compliance with § 29.685(d)(1) and (4), and § 29.685(f).
(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in § 29.397, that part of the system may be substantiated for lower loads as prescribed.
(i) The limit design loads are those loads specified in § 29.397;
The limit design loads are the maximum that can be obtained in normal operation, including any single power boost system failure, except for objectives stated for a rugged system; and
(iii) In lieu of a rational analysis, the limit design loads may be 0.60 of the loads specified in § 29.397.
(iv) For example, if a control surface servo tab or a small elevator is a part of the rotorcraft design, the control system for this part must be stiff enough to react the control surface loads without failure and to provide enough surface deflection to control the rotorcraft. These limit loads may be 60 pounds fore and aft and 40 pounds laterally on the cyclic control stick in lieu of a rational analysis and may be the maximum loads that can be obtained in normal operation.
If a hydraulic power actuation or boost system is part of the rotorcraft design, the design limit load for the affected parts of the control system will be the maximum output force of the boost at normal operating pressure added to the limit design loads resulting from the loads specified in § 29.397. If a single failure in the power portion of the hydraulic system results in actuator forces that exceed the 262 Par 153
maximum output force at normal operating pressure, the highest output loads must be used as noted in Subparagraph (3) (Il). This hydraulic system failure standard is specified in § 29.695(a)(1) as well.
(4) Controls proof and operation test is required by §$ 29.307 (b), 29.681, and
29.683. This test is conducted using the design limit loads approved under § 29.395(b).
See Paragraphs 282 and 283 of this document.
153A. $ 29.395 (Amendment 29-30) CONTROL SYSTEM.
a. Explanation. Amendment 29-30 clarifies that the loads in § 29.395(b) apply to power "control" systems not just power "boost" systems; and the limit pilot forces prescribed in § 29.397 are required to be applied in conjunction with the forces from normally energized power devices. The amendment may increase required loads for systems if operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. If so, the system is required to withstand 100 percent of limit pilot forces specified in § 29.397, rather than 60 percent of the limit pilot forces as specified previously.
b. Procedures. The procedures of Paragraph 153 continue to apply except that the increased loads in new Paragraph § 29.355(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction.
154. § 29.397 (Amendment 29-12) LIMIT PILOT FORCES AND TORQUES.
a. Explanation. Design forces are contained in the rule.
(1) Primary controls, pilot and copilot, must be designed for the limit pilot forces specified in Paragraph (a) of the rule.
(2) For other operating controls, such as flap, tab, stabilizer, rotor brake, and landing gear, design limit forces are specified in Paragraph (b).
b. Procedures.
(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis of the control systems submitted in compliance with other rules.
(2) Operation tests of the control systems noted in other rules require application of these forces also.
Par 153 263
155. § 29.399 DUAL CONTROL SYSTEM.
a. Explanation. Design limit loads are specified for dual control systems. Pilot effort forces applied in opposition and in the same direction are required for dual control systems.
b. Procedures.
(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis submitted for compliance with the other rules.
(2) Operation tests of the control systems, noted in other rules, require application of these forces also.
156. § 29.401 (Amendment 29-4) AUXILIARY ROTOR ASSEMBLIES.
a. Explanation.
(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 29.923 and structural strength substantiation is required. Section 29.401(b) specifically refers to structural strength substantiation for centrifugal loads resulting from maximum design rotor RPM. Due to the pitch feathering requirements, auxiliary rotors typically have detachable blades.
(2) The rotor blade structure must have sufficient strength to withstand not only aerodynamic loads generated on the blade surface, but also inertial loads arising from centrifugal, coriolis, gyroscopic, and vibratory effects produced by this blade movement.
Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to assure that the blades will maintain the desired aerodynamic characteristics. As a design objective, the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads that may increase the operating stress levels. Blade weight and strength should be optimized. Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by § 29.571.
(3) Section 29.1509 defines the design rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits.
b. Procedures.
(1) The endurance tests prescribed by §$ 29.923 and 29.927 require achieving certain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests.
264 Par 155