not more than 20 percent to maximum collective or alternately to the collective setting required to become partially airborne (when the cyclic is displaced as noted).
(7) Typically, articulated rotor aircraft have natural frequencies on the blade in lag of approximately 0.3 times the power-on main rotor RPM; soft in-plane rotors have natural frequencies approximately 0.7 times the main rotor RPM. Therefore, for example, for a rotorcraft with an in-plane frequency of 0.3/rev, operating at 300 RPM, and with 6 inches of total lateral cyclic stick displacement, the stick should be rotated for 5 revolutions in a 0.6-inch diameter circle at ((1-.03) x 300 RPM) or 3.5 cycles per second to attempt excitation of possible resonant frequencies. At the conclusion of the excitation, the cyclic stick should be returned to the neutral position while continuing the recording of data listed in paragraph b(13).
(8) The complete program should again be repeated with cyclic excitation inputs from the directional and longitudinal controls, if critical for the type of rotorcraft being evaluated.
(9) If onset of ground resonance is encountered, the typical recommended corrective action is to increase the collective pitch and rotor speed and become airborne. However, lowering the collective pitch has been effective for some designs and is considered a satisfactory procedure if resonance can be consistently avoided.
(10) Landings should be made at the maximum touchdown speed proposed with the rotor speed stabilized.
(11) Special Considerations:
(i) The influence of variables including environmental effects, corresponding aircraft component characteristic changes, operational parameters, and surface conditions should be investigated over the ranges proposed for certification.
Additionally, the potential of misservicing and possible failure modes should be evaluated. For ground resonance qualification, where practical, variations from the baseline test configuration may be accomplished by either ground run (§ 29.663(b)) requires investigation of probable ranges of damping), analyses, component tests, aircraft shake test, the specification of special operational procedures in the rotorcraft flight manual, or combination thereof. Detailed and rational analyses showing acceptable correlation to the baseline tests, and for which the input parameters were verified by drawings, calculations, component static or dynamic tests, or by aircraft shake tests simulating the conditions/configurations in question, may be used to limit testing to only those variables and operational conditions showing marginal or unacceptable system damping. All operational limitations should be clearly stated in the rotorcraft flight manual. A report of the analytical and/or test results should be permitted per § 29.663.
(ii) Potential instability while airborne, called “air resonance” may occur due to the dynamic coupling of the rotor flexibility and the pylon restraint flexibility. The Page B - 123
same considerations apply to air resonance as to ground resonance except that the pylon restraint variables replace the landing gear variables. Air resonance should be addressed in the certification program.
(iii) When operating on the ground, there may be a tendency for the aircraft to exhibit a “ground bounce.” For many configurations, this is a benign, although undesirable phenomenon which may be aggravated by pilot induced oscillations (PIO), particularly if there is little or no friction on the collective.
(12) On rotorcraft with fully articulated rotor heads equipped with landing gear oleos in either skid or wheel configuration, there are tendencies for ground bounce to occur when light on the oleos, either just prior to takeoff or just after landing contact, or during a power assurance check. This bounce may induce ground resonance, particularly if the intensity of the bounce is aggravated by PIO. The corrective action is either to lift off to a hover or to positively lower the collective and remain on the ground.
(13) Instrumentation and Data Acquisition.
(i) Atmospheric Conditions (to be manually noted):
Altitude OAT Wind Velocity
(ii) Aircraft Configuration (to be manually noted):
Gross Weight C.G.
Tire Pressure Landing Gear Oleo Pressure
(iii) Instrumentation (for recording during test).
Main Rotor RPM.
Time history of cyclic control fore-and-aft and lateral stick position Time history of collective control stick position Time history of rotor damper motion* Time history of pylon component motion* Time history of landing gear (oleo) motion* Time history of aircraft motions* *As required to obtain modal damping Page B - 124
SUBPART B - FLIGHT
MISCELLANEOUS FLIGHT REQUIREMENTS AC 29.251. § 29.251 VIBRATION.
a. Explanation.
(1) Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition (rule statement).
(2) This flight requirement may be both a qualitative and quantitative flight evaluation. Section 29.571(a) contains the flight load survey requirement that results in accumulation of vibration quantitative data. Section 29.629 generally requires quantitative data to show freedom from flutter for each part of the rotorcraft including control or stabilizing surfaces and rotors. See paragraphs AC 29.571 and 29.629 for these two rules.
(3) Review Case No. 70 (reference FAA Order 8110.6) contains a policy 0statement concerning compliance with this rule. This policy statement is condensed here for convenience:
“The rotorcraft must be capable of attaining a 30° bank angle (turn), at VNE, with maximum continuous power (maximum continuous torque) without encountering excessive roughness/vibration. The FAA/AUTHORITY requires the maneuver demonstration to provide the pilot with some maneuver capability at VNE , and further to provide the pilot some margin away from roughness when operating in turbulence.”
(This maneuver may result in a descent or a climb.)
(4) Section 29.1505 pertains to VNE determination. Section 29.1509 pertains to rotor speed limits determination. See paragraphs AC 29.1505 and AC 29.1509.
b. Procedures.
(1) During the company flight test program, the rotorcraft is flown to the appropriate rotor and airspeed limits at several weights to prove that the rotorcraft is free from excessive vibration under appropriate speed, power, and weight conditions.
The flight loads survey quantitative data (reference § 29.571) and the applicant’s qualitative and quantitative flight test data must also prove compliance with the requirement prior to issuing an authorization for official FAA/AUTHORITY flight tests.
(2) The flight load survey data obtained under § 29.571(a) will contain measured data concerning proof of freedom from flutter and excessive vibration.
Pertinent critical flight conditions will be reinvestigated during FAA/AUTHORITY flight tests. The specific condition or conditions necessary to demonstrate compliance with § 29.251 varies with the rotorcraft design, and with the minimum and maximum rotor Page B - 125
speeds, VNE and VD speeds, and weight and CG position. An illustration of the speed and RPM demonstration is shown in figure AC 29.251-1. Also see subparagraph b(4).
(3) The airspeed and rotor speed limits investigated and established under §§ 29.33, 29.1503, 29.1505, and 29.1509 are also investigated and made a matter of record in the flight loads survey data. During the official FAA/AUTHORITY/TIA flight tests, critical parts of the rotorcraft may have limited instrumentation to reinvestigate and confirm that the critical conditions investigated during the flight load survey are satisfactory and do not result in excessive vibration. Use of instrumentation is optional if the flight loads data (reference paragraph AC 29.571) are conclusive.
(4) FAA policy for certification (Review Case No. 70) requires a “rotor roughness” flight demonstration of a 30° bank angle left and right, at maximum continuous power (MCP) (maximum continuous torque which may be in excess of the maximum continuous temperature limit), at VNE. To provide the pilot with some margin from roughness, the FAA requires maneuver demonstrations of 30°banked turns at VNE without encountering excessive roughness. The maneuver should be conducted with the rotor speed at the minimum RPM and maximum RPM limits. During the flight load survey, this condition should be investigated and data recorded to assure hazardous loads are not encountered for this “unusual” condition. As indicated, the flight condition will be reinvestigated during the FAA/AUTHORITY flight tests. See paragraph b(2) for illustration of this speed and RPM demonstration.
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0° Bank, Autorotation Level Forward Flight, 105% Max RPM Power-Off g 4 Power-On 30° Bank, Power-On Max RPM Power-Off 02 Max RPM Power-On a7 д 6 ROTOR RPM -Min RPM Power-On ш 8 д 5 Min RPM Power-Off 01 95% Min RPM Power-Off o 3 VNE 1.11 VNE
AIRSPEED
1. Autorotation at 1.11 VNE (AR) at minimum placard rotor speed.
2. Autorotation at 1.11 NNE (AR) at maximum placard rotor speed.
3. Autorotation at NNE (AR) at power-off minimum design limit rotor speed
4. Autorotation at NNE (AR) at power-off maximum design limit rotor speed
5. Forward flight 1.11 VE at minimum power-on rotor speed.
6. Forward flight 1.11 VNE at maximum power-on rotor speed.
7. Right and left turn at VE at maximum power-on rotor speed with 30° bank angle.
8. Right and left turn at VE at minimum power-on rotor speed with 30° bank angle.
Note: VNE (AR) may be less than VNE • FIGURE AC 29.251-1 DEMONSTRATION POINTS Page B - 127
CHAPTER 2. PART 29
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
SUBPART C - STRENGTH REQUIREMENTS
GENERAL AC 29.301. § 29.301 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 must be distributed appropriately or conservatively and significant 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.
AC 29.303. § 29.303 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.
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AC 29.305. § 29.305 STRENGTH AND DEFORMATION.
a. Explanation.
(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.
(2) If static tests are used to show compliance with this rule, the structure must support ultimate loads for 3 seconds without failure. Alternatively, dynamic tests simulating actual load applications may be used.
(3) Section 29.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See paragraph AC 29.307.
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.
(2) Any test results reports must contain data and information showing the test results comply with the standard.
(3) When dynamic tests are not used to substantiate the ultimate strength of structure subject to significant dynamic response under load, the analytical substantiation should consider flexibility effects and rate of load application (tail boom strength under landing loads is an example of a strength which needs dynamic amplification effects considered).
AC 29.307. § 29.307 (Amendment 29-4) 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 § 29.307(b)(6).
(2) “Structural analysis may be used only if the structure conforms to those for which experience has shown the structural analysis method to be reliable.”
(3) Fatigue substantiation requirements are explained further in paragraph AC 29.571.
b. Procedures.
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(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 and dynamics (vibration and stability) reports and fatigue, static, dynamic, 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 accepted by the FAA/AUTHORITY for the test. Test fixtures and instrumentation must 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 is not intended or envisaged by this policy.
Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to assure 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 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.
(4) Section 29.307(b) specifies certain tests. Test proposals must be approved prior to conducting official FAA/AUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests.
AC 29.307A. § 29.307 (Amendment 29-30) PROOF OF STRUCTURE.
a. Explanation. Amendment 29- 30 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 material 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 Page C - 3
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-17B, Vol I, Rev. 1E; Vol. II, Rev.
D; Vol. III, Rev. E (or later versions) are acceptable sources of data and procedures to show compliance with environmental effects of metallic and composite materials, respectively.
AC 29.309. § 29.309 DESIGN LIMITATIONS.
a. Explanation.
(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant must 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 29.321. § 29.321 GENERAL - FLIGHT LOADS.
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 § 29.337 through § 29.351. This rule requires flight load det ermination 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 anal ysis data must comply with the rule.
AC 29.337. § 29.337 (Amendment 29-30) 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 weights.
(4) Load factors exceeding these “minimums” may be used.
b. Procedures.
(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 Page C - 5
AC 29-2C, Chg 4 5/1/2014 used, an analytical study and flight demonstration are required. Maximum available rotor lift with both power on and power off must be considered when substantiating maneuver load factors less than the specified values.
(i) The maximum positive design load factor is +3.5 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 29.337-1. 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 § 29.337(b)(2) since the design load factor varies with gross weight.
(ii) 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 § 29.337(b), the applicant must conduct the flight demonstration required by § 29.337(b)(1). The flight test personnel must determine that the demonstration is conducted in a manner to show that 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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