138. § 27.339 (Amendment 27-11) RESULTANT LIMIT MANEUVERING LOADS.
a. Explanation. The rule specifies or defines the application of rotor and lift surtace loads to the rotorcraft.
(1) The design maneuvering load factors required by § 27.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.
(2) Sections 27.321, 27.337, and 27.341 all complement one another and result in the derivation of design flight loads that will be imposed to ensure structural integrity of the rotorcraft.
(3) The following assumptions and conditions are specified in the rule.
(i) The rule requires application of appropriate loads at each rotor hub and auxiliary lifting surface.
(ii) Power-on and power-off flight with maximum design rotor tip speed ratio and specific conditions that must be considered.
(Il) Rotor tip speed ratio, defined in the rule, has been carried forward from the initial rotorcraft certification rules issued in 1946. The rotor tip speed ratio is a basic parameter used in calculating rotor aerodynamic forces.
b. Procedures.
(1) The rule specifies an acceptable assumption concerning application of the rotorcraft maneuvering loads.
(2) The rotor tip speed ratio is a parameter found in textbooks and other books such as NACA Report No. 716. The equation in the rule contains angle "a." Report No. 716 also defines angle "a" as the angle of attack of the rotor disk. This definition is more easily understood than the definition contained in the rule.
(3) The rotorcraft design loads are derived as prescribed by $$ 27.321, 27.337, and 27.341. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.
139. §27.341 GUST LOADS.
a. Explanation.
(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical gusts from hovering to 1.11 VNE (i.e., VD).
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AC 27-1A 7130/97
(2) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward.
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.
(ii) 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 is generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.
(Hi) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.
(iv) The gust velocity may be assumed uniform across the rotorcraft.
(V) 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 ensure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load factor.
(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA -Ames Specialist's Meeting on Rotorcraft Dynamics, February 13-15, 1974. The paper, entitled, "Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects,i was written by P.J.
Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.
140. § 27.351 (Amendment 27-26) YAWING CONDITIONS.
a. Explanation. The standard was added by Amendment 27-26. It requires proof of a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope 186 Par 139
should cover minimum forward speed or hover to the lesser of VH or VNE for "power-on" condition, not "power-off" since Vy is a part of the standard. The rotorcraft should be structurally safe for the thrust capability of the directional control system as stated.
(1) The rotorcraft structure should 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 derived from § 27.397(a)), is required for the conditions cited in the standard. A control system rate limiter or a yaw damper may be used as part of the type design required equipment, if elected. The total displacement is therefore a function of time as well as the maximum effort applied such as 130 pounds.
(i> At low airspeeds from 0 to 0.6 VNE, 90° yaw (sideward flight) is specified as the design limit.
(Ii) At high airspeeds VH or VNE), stabilized yaw angle (stabilized sideslip) or 15° sideslip, whichever is less, is specified to be substantiated.
(iii) 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 § 27.351(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.
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 FH1100 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.
(ii) 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.
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NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (§ 27.351(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 § 27.351(b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 27.351(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 § 27.351(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 logical assessment of the various yaw conditions may be used to reduce the load derivation and analysis workload for critical rotorcraft design conditions.
(vi) The rotorcraft structure should be analyzed and/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 and 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 0.6 VNE) and 15° yaw when they are applied at the lesser of VNE or VH, with a straight line variation from 0.6 VNE to VH or VNE. The rotorcraft does not need to be capable of attaining the 90° and 15° yaw. They should be considered as the maximum sideslip limits.
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90°
YAW DEGREES (LONGITUDINAL AXIS) → 15° MAX.
0°
0.6 VNE AIRSPEED
LESSER OF
VA OR VNE FIGURE 140-1. SAMPLE YAW/FORWARD SPEED DIAGRAM Par 140 189
(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. However, this test is not a requirement.
(2) FAR § 27.351(b)(1) incorrectly references § 27.395(a) for the maximum pilot forces. The correct reference should be § 27.397(a).
141. § 27.361 (Amendment 27-23) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft shall be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. Amendment 27-23 separated the standard into paragraphs for turbine and reciprocating engine limit torque values.
(2) Turbine engine limit torque for design purposes (Amendment 27-23) was redefined into four cases and the torque values determined will be used. For example, sudden engine stoppage is introduced as one of the cases which is applied to the engine and the engine suspension and restraint system. Emergency operation of governor-controlled turboshaft engines is another case.
(3) Torque factors are also specified for reciprocating engines having two or more cylinders in Paragraph (b) of the standard.
(4) Sections 27.547(e)(1)(ii) and 27.549(d), respectively, refer to the application of engine torque to design of main rotor structure and engine mount and adjacent structure.
b. Procedures.
(1) The engine torque associated with the maximum continuous (MC) power condition for reciprocating engines should be multiplied by the appropriate torque factor to obtain the limit engine torque value used for structural substantiation of the rotorcraft.
(2) The torque values associated with MC power at 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.
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(3) Turbine engine limit torque values are determined for the four cases specified. Two cases are related to the endurance test of §§ 27.923 and 27.927.
(4) For sudden stoppage of turbine engines the engine manufacturers can reasonably provide FAA/AUTHORITY approved data to the applicant on inertia of rotating parts and the deceleration time expected in the event of sudden engine stoppage. This condition usually generates critical loads in the engine mounting and restraint system. These manufacturer's data should be acceptable for use in compliance with this part of the standard.
142.-151. RESERVED Par 141 191 (thru 201)
SECTION 9, CONTROL SURFACE AND SYSTEM LOADS
152. § 27.391 GENERAL.
a. Explanation. This general standard 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 shall contain the loads dictated by the referenced rules. (See Paragraphs 153, 154, 155, 156, 157, 158, and 159 of this document.)
152A. § 27.391 (Amendment 27-26) GENERAL.
a. Explanation. Amendment 27-26 adds an explicit reference to § 27.427, Unsymmetrical Loads (AC Paragraph 160), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 27.399, Dual Control System (AC Paragraph 155), is also added for clarification. In addition, §§ 27.401, 27.403, and 27.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.
b. Procedures. The referenced AC paragraphs become 153, 154, 155, 158, and
160.
153. § 27.395 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 minimum design loads specified in §§ 27.397 and 27.399 (for dual control systems).
Except reduced design loads, not less than 0.60 of those specified in §§ 27.397 and
27.399 for dual control system, may be used as specified. The standard also applies to those control systems that may have more than one stop in a system. The design loads must be imposed on the system from the pilot's control to any stop in the control system.
(2) Minimum design loads imposed on the control system from a stop to a rotor blade or a control surface or device shall be:
(i) The maximum pilot forces obtainable in normal operation; and If low operational loads may be exceeded as noted in § 27.395(b)(2), the system shall support without yielding 0.60 of the loads specified in §§ 27.397 and
27.399 for dual control systems.
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(3) Section 27.695 concerns standards for a power boost and power-operated control system. This standard, in effect, imposes a fail-safe standard for hydraulic aspects of a control system. Where appropriate to a particular design, the control system must therefore sustain without yielding, the maximum output force of the actuator when complying with § 27.395(a). The pilot input forces are not added to the actuator output forces according to this standard for normal category rotorcraft. These forces are independently applied to the control system.
(4) Control system design features and tests requirements are found in §§ 27.619 and 27.625, respectively. Special factors such as casting, bearing, and fitting factors that may be appropriate for the design are contained in §§ 27.619 and
27.625, respectively.
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.
A static test proposal for testing the control system to show compliance with the rules should be approved before conducting the test. Where compliance is to be determined by tests, limit load tests, as discussed in Paragraph 284 of this document, and/or ultimate load tests may be performed. Test results shall be documented.
(ii) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 27.623 and 27.625 must be used as specified.
Tests may not be required when adequate similarity of systems and support structure is determined and where adequate structural analysis is furnished.
(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in §§ 27.397 and 27.399, that part of the system may be substantiated for lower loads as prescribed.
(i) The limit design loads are those loads specified in §§ 27.397 and
27.399;
The maximum that can be obtained in normal operation and that is allowed by the system; except (її) The limit design loads may not be less than 0.60 of the limit pilot forces specified.
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(iv) For example, if a small control surface or servo tab is lightly loaded, its control system must be stiff enough to react the control surface loads and to provide surface deflection to control the rotorcraft. The normal operational loads may be very low, such as 10 pounds maximum. Nonetheless, the design limit load shall be
0.60 times the limit single pilot forces specified in § 27.397. Note that the system must not yield under these loads.
(v) For example, if a dual but primary manual control system such as a tail rotor control is lightly loaded, the control system, from the stops to the rotor blades, may be designed for minimum loads equal to 0.60 times the limit dual pilot forces specified in § 27.399.
(vi) If a power actuator is a part of a rotor control system, the design limit force for the affected parts shall be the maximum output force of the actuator at any operational condition (including any load/pressure after a single failure in the hydraulic system).
(4) Controls proof and operation test is required by §§ 27.307(b)(2) and (b)(3),
27.681, and 27.683. This test is conducted using the design limit loads approved under § 27.395. (See Paragraphs 284 and 285 of this document.)
153A. § 27.395 (Amendment 27-26) CONTROL SYSTEM.
a. Explanation. Amendment 27-26 extensively rewrites § 27.395(b) to more clearly incorporate design condition loads for typical powered control systems. New requirements include substantiation for loads resulting from "each normally energized power device, including any single power boost or power activator failure in the control system. There are also new minimum loads for control system designs in which operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. The old loads were 0.60 times the limit pilot forces of § 27.397; the new loads are 100 percent of the limit pilot forces specified in § 27.397.
b. Procedures. The procedures of Paragraph 153 continue to apply except that the increased loads in § 27.395(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. § 27.397 (Amendment 27-11) LIMIT PILOT FORCES AND TORQUES.
a. Explanation. Design forces are contained in the rule.
(1) Primary controls, pilot and copilot, should be designed for the limit pilot forces specified in Paragraph (a) of the rule unless higher torces are used.
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(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 may 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.
155. § 27.399 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 may 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. § 27.401 (Amendment 27-3) AUXILIARY ROTOR ASSEMBLIES.
a. Explanation.
(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 27.923, and structural strength substantiation is required. Section 27.401(b) specifically refers to structural strength substantiation of detachable blade systems for centrifugal loads resulting from maximum design rotor RPM.
(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 ensure 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 Par 154 205
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 § 27.571.
(3) Section 27.1509 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 §§ 27.923 and 27.927 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.
(2) Structural substantiation of the auxiliary (tail) rotor is required to ensure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by § 27.927 (b).
(•) The maximum and minimum operating rotor speed, power-off, is 95 percent of the maximum design speed and 105 percent of the minimum design speed, respectively.
The rotor operating speed limits shown during the official FAA/AUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds.
(iii) The auxiliary rotor generally has a positive and negative pitch limit that ensures adequate directional control throughout the operating range of the rotorcraft. The power-off rotor speed limits are generally broader than the power-on rotor speed limits because of the required autorotational rotor speed characteristics.
Thus, the auxiliary rotor design conditions concern the maximum and minimum design rotor speeds in conjunction with the maximum positive or negative pitch thrust, as appropriate. Thrust capability and precone angle of the rotor, if any, will significantly influence the rotor design loads. The variations in rotor design features and an example of substantiation would be too lengthy to include here. However, ANC-9, "Aircraft Propeller Handbook" contains principles that may be applied to tail rotor designs. Tail rotors may be considered a special propeller design.
(iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearing manufacturer's ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of §§ 27.923 and 27.927 is a means of proving that use of a particular bearing is satisfactory.
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