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CHAPTER 2. PART 27
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
SUBPART C - STRENGTH REQUIREMENTS
STRENGTH REQUIREMENTS - GENERAL AC 27.301. § 27.301 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 are specified to be distributed appropriately or conservatively and significant changes in distribution of the loads, as a result of deflection, are specified to be taken into account.
b. Procedures. The design criteria report and/or design loads report should contain data that comply with the rule.
AC 27.303. § 27.303 FACTOR OF SAFETY.
a. Explanation.
(1) Unless otherwise provided by Part 27, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will ensure that the design strength of the rotorcraft is greater than the design loads contained in Part 27.
(2) Other rules, §§ 27.561(b)(3) and 27.787(c), specify use of defined ultimate inertial forces for protection of occupants.
b. Procedures.
(1) The design criteria report and/or design loads report should contain data that include the appropriate factor of safety.
(2) The factor of safety multiplies the limit external and inertial loads. The rule does allow the application of this factor to the resulting "limit internal" stresses if it is more conservative.
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AC 27.305. § 27.305 STRENGTH AND DEFORMATION.
a. Explanation.
(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.
(2) The structure is required to be able to support, in a static test, ultimate loads for 3 seconds without failure, or dynamic tests simulating actual load application may be used.
(3) Section 27.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See paragraph AC 27.307.
b. Procedures. Any test results, static or dynamic, should satisty 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 Part 27 should contain the criteria stated in the rule.
(2) Any test results reports shall contain data and information showing the test results comply with the standard.
AC 27.307.§ 27.307 (Amendment 27-3) PROOF OF STRUCTURE.
a. Explanation.
(1) The rule requires compliance with the strength and deformation requirements 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 § 27.307 (b)(6).
(2) Structural analysis rather than load tests may be used only if the structure conforms to those for which experience has shown the structural analysis method to be reliable.
b. Procedures.
(1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the critical loading conditions will be selected for analytical substantiation in structural (static and fatigue) reports, dynamics (vibration and stability) reports, and in fatigue, static, dynamic, or operational test reports.
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(2) Whenever tests are used or required, a test proposal or plan should be approved prior to the tests. The test article should have received conformity inspections and should have been accepted by the FAA/AUTHORITY for the test. Test fixtures and instrumentation should also be acceptable to the FAA/AUTHORITY (using DERS as appropriate) prior to the start of the test. The quality control office of the applicant or other qualified personnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA/AUTHORITY 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 are not intended or envisaged by this policy.
Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to ensure correct location, direction, and magnitude of loads and other critical test parameters.
(3) Structural analysis has been accepted for rotorcraft in place of static tests.
Generally, the rotorcraft airframe should have natural frequencies remote to predominant rotor excitation sources, including higher harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this analytical data to verify accuracy.
(4) Paragraph (b) of the rule specifies certain tests. Test proposals should be approved prior to conducting official FAA/AUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests.
AC 27.307A. § 27.307 (Amendment 27-26) PROOF OF STRUCTURE.
a. Explanation. Amendment 27-26 adds the requirement to account for the environment to which the structure will be exposed in operation. This change is intended to codify recent FAA/AUTHORITY and industry practices for the consideration of environmental effects in showing "proof of structure."
b. Procedures. All of the policy materials pertaining to this section remains in effect with the following additions:
(1) For either tests or an analysis, environmental effects are now explicitly required. Consideration of loss of strength and stiffness of metals with elevated temperatures and loss of strength and stiffness of composite materials from exposure to heat, moisture, or other operational environments is now required and should be documented in analyses and test reports.
(2) MIL-HDBK-5F; AC 20-107B; or MIL-HDBK-17, Vol. I, Rev 1E; Vol. II, Rev. D; Vol. III, Rev E (or later versions) are acceptable sources of data and procedures Page C - 1
to show compliance with environmental effects of metallic and composite materials, respectively.
AC 27.309. § 27.309 DESIGN LIMITATIONS.
a. Explanation.
(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant is required to establish these structural limitations to facilitate design of the rotorcraft.
(2) Refer to the rule for the specific 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.
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SUBPART C - STRENGTH REQUIREMENTS
FLIGHT LOADS AC 27.321 § 27.321 (Amendment 27-11) FLIGHT LOADS--GENERAL.
a. Explanation.
(1) The rule specifies the way the loads will be applied to the rotorcraft. It requires load analysis from minimum to maximum design weight. Any practical distribution of disposable loads must be included in the analysis.
(2) Paragraph (a) of the rule states: "The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft, and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravity."
b. Procedures.
(1) Derivation of the flight loads is required by and specified in §§ 27.337 through 27.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads.
(2) The application of the design loads derived from the flight load factor will be as specified. The flight loads analysis data must comply with the rule.
AC 27.337 § 27.337 (Amendment 27-26) LIMIT MANEUVERING LOAD FACTOR.
a. Explanation. The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors.
(1) A range of design positive load factors from +3.5 to +2.0 may be used.
(2) A range of design negative load factors from -1.0 to -0.5 may be used.
(3) Load factors inside the range of +3.5 to -1.0 may be used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote and the selected load factors are appropriate to each weight condition between design maximum and minimum weight.
(4) Load factors exceeding these "minimums" may be used.
b. Procedures.
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AC 27-1B, AC 27-1B, Chg 4 3/31/2014
(1) The applicant may elect to substantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is used, an analytical study and flight demonstration are required.
(i) The maximum positive design load factor of +3.5 is generally at a weight below maximum gross weight. The maximum thrust capability of the main rotor, combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor-gross weight curve is shown in figure AC 27.337-1. Note the minimum positive design load factor is +2.0 even though the required analysis and flight demonstration may prove the rotorcraft is not capable of achieving this load factor. This curve also illustrates compliance with § 27.321(b)(1) since the design load factor varies with gross weight.
(її) The largest negative design load factor is -1.0; however, several current rotorcraft designs are not capable of achieving a negative load factor.
Therefore, -0.5 has been an acceptable structural design negative load factor for certain rotorcraft designs.
(2) Whenever the applicant analytically substantiates the lower load factors allowed by § 27.337(b), the flight demonstration required by § 27.337(b) must be conducted. The flight test personnel should determine that the demonstration shows the probability of exceeding the selected design load factors (those factors less than +3.5 and more than -1.0) is extremely remote.
(3) A numerical value has not been assigned to "extremely remote" in this standard.
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- Fepresents curve ol Design thrust + lift Load Factor —#ŽÛ
- 1.0 Minimum Gross Weight Lbs. Махітціт
LOAD FACTOR GROSS WEIGHT CURVE FIGURE 27.337-1 Page C - 7
AC 27.339 § 27.339 (Amendment 27-11) RESULTANT LIMIT MANEUVERING LOADS.
a. Explanation. The rule specifies or defines the application of rotor and lift surface 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 surtace 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.
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.
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.
AC 27.341 $ 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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(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.
(iii) 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 AC 27.413.
(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
(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," was written by P.J.
Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helptul also.
AC 27.351. § 27.351 (Amendment 27-26) YAWING CONDITIONS.
a. Definitions.
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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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