AC 27-1B, Chg 2 4/25/06
(1) Suddenly. For the purpose of this section, 'suddenly' is defined as an interval not to exceed 0.2 seconds for complete control input. A rational analysis may be used to substantiate an alternative value.
(2) Zero Yaw. Normal, 1-g, level flight condition with either zero bank angle or zero sideslip.
b. Explanation. The rule requires a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed, or hover, to VH 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 yawing conditions. The standard does not require a structural flight demonstration. It is a structural design standard.
(2) This standard applies only to power-on conditions. Autorotations need not be considered.
(3) This standard requires the maximum allowable rotor RPM consistent with the flight conditions, including special operational rotor settings.
(4) For the purposes of this section, the analysis may be performed at international standard atmosphere (ISA) sea level conditions.
(5) The rotorcraft structure must be designed to withstand the loads for the specified sideslip conditions. This includes, but is not limited to:
(i) Main cabin, tailboom, and vertical control surfaces.
(ii) Tail rotor structures, including the fitting attachments to the frame.
(iii) Windows, doors, and other transparencies.
(iv) Landing gear and retracting mechanism.
(v) Fairings and cowlings.
(6) Maximum displacement of the directional control, except as limited by pilot effort (§ 27.397(a)), is required for the conditions cited in the rule.
Control system limiting devices may be used, however the probability of failure or malfunction of these systems) should be considered (see Figure AC 27.351-2).
This evaluation may include Flight Manual Limitations, if failure of the system is reliably indicated to the crew.
(7) Both right and left yaw conditions should be evaluated.
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(8) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust (superpositioning).
(9) Loads associated with sideslip angles exceeding the values of Figure AC 27.351-1 do not need to be considered. The corresponding points of the maneuver may be deleted.
c. Procedure. The design loads should be evaluated within the limits of Figure AC 27.351-1 or the maximum capability of the rotorcraft, whichever is less; at speeds from zero to VH or VNE, whichever is less, for the following phases of the maneuver:
(1) With the rotorcraft at an initial trim condition (1 g level flight and zero yaw), the cockpit directional control is suddenly displaced to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 27.397(a). This is intended to generate a high tail rotor thrust.
(2) While maintaining maximum cockpit directional control deflection, within the limitation specified in c(1) of this AC paragraph, allow the rotorcraft to yaw to the maximum transient sideslip angle or to the value defined in Figure AC 27.351-1, whichever is less. This is intended to generate high aerodynamic loads.
(3) Allow the rotorcraft to stabilize at the maximum steady-state sideslip angle.
In the event that the maximum steady state angle is greater than the value defined in Figure AC 27.351-1, the rotorcraft should be trimmed to the value of the angle using less than maximum cockpit directional control deflection.
(4) With the rotorcraft yawed to the static equilibrium sideslip angle specified in c(3) of this AC paragraph, the cockpit control is suddenly returned to its initial trim position. This is intended to combine a high tail rotor thrust and high aerodynamic restoring torces.
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AC 27-1B, Chg 2 4/25/06
SIDESLIP ANGLE (deg. ) 90 15 →
0.6 Vne Vh or Vne whichever is less
AIRSPEED Figure AC 27.351-1 Safety Factors for Probability of Failure Page C - 12
• For static strength substantiation, each part of the structure should be able to withstand without failure, the loads generated by the maneuver described in the rule multiplied by a factor of safety depending on the probability of being in this failure state. The factor of safety is defined in the figure below:
10:9 van - Probabilly or belng in tallura condition T Qj= (Tj)(Pj) where:
Tj = Average flight time spent with a failed control limiting system j (in hours Pj = Probability of occurrence of failure of the control limiting system j (per hour) Note: If Pj is greater than 10-3 per flight hour then a 1.5 factor of safety should be applied to all limit load conditions specified in this standard.
Figure AC 27.351-2 Page C - 13
AC 27.361 § 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.
(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.
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SUBPART C - STRENGTH REQUIREMENTS
CONTROL SURFACE AND SYSTEM LOADS AC 27.391. § 27.391 CONTROL SURFACE AND SYSTEM LOADS--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 AC 27.395, AC 27.397, AC 27.399, AC 27.401, AC 27.403, AC 27.411, and AC 27.413.)
AC 27.391A. § 27.391 (Amendment 27-26) GENERAL.
a. Explanation. Amendment 27-26 adds an explicit reference to § 27.427, Unsymmetrical Loads (paragraph AC 27.427), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 27.399, Dual Control System (paragraph AC 27.399), 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 paragraphs AC 27.395, AC 27.397, AC 27.399, AC 27.411, and AC 27.427.
AC 27.395. $ 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:
The maximum pilot forces obtainable in normal operation; and Page C - 15
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.
(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 AC 27.307.
(i) 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 AC 27.681, 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 Page C - 16
(¡¡i) The limit design loads may not be less than 0.60 of the limit pilot forces specified.
(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 AC 27.681 and AC 27.683.)
AC 27.395A. § 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 AC 27.395 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.
AC 27.397. § 27.397 (Amendment 27-11) LIMIT PILOT FORCES AND TORQUES.
a. Explanation. Design forces are contained in the rule.
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(1) Primary controls, pilot and copilot, should be designed for the limit pilot forces specified in paragraph (a) of the rule unless higher forces are used.
(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.
AC 27.399. $ 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.
AC 27.401. $ 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 Page C - 18
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 § 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).
(i) 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.
(її) 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.
(m) 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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(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by §§ 27.619, 27.621, 27.623, and 27.625, respectively. The fitting factor of 1.15 must be applied in the analysis of the tail rotor installation.
AC 27.401A. § 27.401 (Amendment 27-27) AUXILIARY ROTOR ASSEMBLIES.
a. Explanation. Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341.
b. Procedures. The policy material pertaining to this section is retained as supplemental information.
AC 27.403. § 27.403 AUXILIARY ROTOR ATTACHMENT STRUCTURE.
a. Explanation.
(1) The auxiliary rotor attachment structure(s), which is considered to include gearboxes, must be designed to withstand design limit loads that occur in flight and on landing. These design loads that generally consist of the following must be established for the particular flight and landing condition under consideration.
Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gearboxes, combined with--
(ii) Thrust and torque loads developed by the auxiliary rotors.
The linear and angular acceleration loads imposed by the weight of the tail rotor and gearbox are generally derived from airframe loads data. Thrust and torque output of the tail rotor are derived during external aerodynamic and landing loads development for pertinent flight and landing conditions.
(2) General rules related to proof of structure loads and factor of safety are §§ 27.307, 27.301, 27.303, and 27.305.
b. Procedures.
(1) The angular and linear acceleration loads combined with appropriate tail rotor thrust and torque for the critical conditions shall be imposed on the tail rotor gearbox mount lugs, the airframe mounting structure, and the attaching hardware.
(2) The yaw and maximum power climb conditions are generally critical.
Landing and maneuvering conditions with and without power may also impose high inertia and rotor thrust and torque loads on the attachment structure.
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(3) The derivation of the loads and conditions is too extensive to include here.
Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, "Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Design."
AC 27.403A. § 27.403 (Amendment 27-27) AUXILIARY ROTOR ATTACHMENT STRUCTURE.
a. Explanation. Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, 27.341.
b. Procedures. The policy material pertaining to this section is retained as supplemental information.
AC 27.411. § 27.411 GROUND CLEARANCE: TAIL ROTOR GUARD.
a. Explanation.
(1) The rule requires specific protection to prevent the tail rotor from contacting the landing surface during a normal landing if it is possible that the tail rotor will contact the surtace. The rule states that it must be impossible for the tail rotor to contact the surface during a normal landing.
(2) If a guard is required, the guard and its supporting structure must withstand suitable design loads.
(3) Section 27.501(c)(1) contains skid landing gear drag requirements that may be applied to the guard design loads.
b. Procedures.
(1) The applicant may submit sketches or drawings showing probable clearance with typical level landing surfaces during normal landings. Typical attitudes such as nose-high autorotation, or autorotation with power-on landing, or other possible tail-low attitudes should be investigated. If the drawings or sketches reveal that it is not likely the tail rotor will contact the landing surface, this minimum clearance with the landing surface may be confirmed during official FAA/AUTHORITY flight tests, such as HV and landing tests. The clearance may be confirmed by having a frangible device of suitable length (i.e., a balsa wood dowel) extending beyond the guard and attached to the tail rotor guard or other appropriate fuselage part. If the device is not damaged, broken, or no contact is made with the surface, compliance has been demonstrated.
(2) If it is possible for the tail rotor guard to contact the landing surface, suitable design loads must be established for the guard. ANC-2a dated March 1948, "ANC Bulletin Ground Loads," paragraph 6.4, entitled "Tail Bumper Criteria," is an acceptable Page C - 21