AC 29.2. § 29.2 (Amendment 29-32) SPECIAL RETROACTIVE REQUIREMENTS.
a. Explanation.
(1) Amendment 29-32 requires a combined shoulder harness and safety belt (also called a torso restraint system) at each occupant’s seat for all rotorcraft manufactured after September 16, 1992.
(2) The design features of the restraint system are mainly contained in this section rather than having to refer to other sections within Part 29 except for a general reference to the differing strength standards between earlier static strength only standards and the static and dynamic strength standards of Amendment 29-29.
(3) Combined safety belt and harness strength standards system follows:
(i) Those rotorcraft type designs certificated to static strength standards alone prior to Amendment 29-29, such as 4 g’s forward may use belt and harness systems, characterized as 1,500 pounds strength systems, provided they comply with those standards. TSO C22f and earlier restraint systems have such ratings. A combined belt and harness with a 1,500 pounds rating, which comply with the Part 29 standards for the rotorcraft type design, but are not necessarily TSO approved, may be approved as a part of the type design. Such design information for a non- TSO’d item would be included in a note on the aircraft type certificate data sheet (TCDS) or specification sheet by part number as “required equipment.” TSO C114- approved torso restraint systems, characterized as 3,000 pounds strength system, may be used provided the design features comply with this section, but no special information on the TCDS is necessary.
(ii) Those rotorcraft type designs certified to dynamic test requirements of Amendment 29-29 should use torso restraint systems approved under TSO C114 or approved under equivalent standards such as those contained in Part 29.
(4) Load Distribution and Design Requirements. Although not stated in § 29.2, a 60 percent and 40 percent load distribution between the safety belt and harness, respectively, is required in § 29.785(g). The safety belt should withstand 100 percent if the safety belt is capable of being used alone. Also, the safety belt or harness attachments to the seat or structure should include the 1.33 factor described in § 29.785(f)(2) of Amendment 29-24 for those rotorcraft with that certification criteria or should include the 1.15 factor as described in § 29.625 (and predecessor § 7.355(c)(2) CAR Part 7) standards for those rotorcraft with the earlier certification criteria. A factor is used whether test results or analysis methods are used for static substantiation of the seating systems. Refer to paragraph AC 29.785b(1)(i) (§ 29.785).
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AC 29-2C, Chg 1 2/12/03
(5) The companion operating rule change of Amendment 91-220, amended § 91.205 (Amendment 91-223), affecting the aircraft equipment requirements.
Operating rule § 91.107(a) already requires use of the harness whenever the aircraft seat is so equipped.
b. Procedures.
(1) A TSO-approved combined safety belt and harness or torso restraint system may be used provided the installation requirements in § 29.2 are satisfied. A combined belt and harness (not necessarily TSO approved) may be approved as a part of the rotorcraft type design and so noted on the aircraft specification or TCDS.
(2) Structural analysis or static test may be used. For those rotorcraft designs that are subject to the dynamic test standards of § 29.562, the torso restraint system is required to be qualified for the particular use or installation in each rotorcraft type design. A dynamic test may be required for alternate restraint systems as well as the originally approved system. TSO C114 approval does not constitute approval for installation of a restraint system in a rotorcraft design subject to dynamic tests.
(i) Paragraph 27.562 of this AC concerns in part the dynamic test standards of Amendment 29-29.
(ii) AC 23-4 dated June 20, 1986, concerns static test procedures for small airplane seats and restraint systems. (Certain small airplanes manufactured after December 12, 1986, should have harnesses for each seat also.) A test proposal for rotorcraft installations may adopt procedures appropriate to the particular installation.
The 60/40 percent distribution is sufficiently achieved when the blocks in Figure 4 of AC 23-4 are used.
(iii) The static design side load for the harness installation may be proven by test or analysis using the load distribution previously noted. For “older” designs, the side load of § 29.561(b)(3)(iii) is 2.0g, and for later designs (Amendment 29-29 and later), it is 8.0g.
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CHAPTER 2. PART 29
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
SUBPART B - FLIGHT
GENERAL AC 29.21. § 29.21 (Amendment 29-24) PROOF OF COMPLIANCE.
a. Explanation.
(1) This section provides a degree of latitude for the FAA/AUTHORITY test team in selecting the combination of tests or inspections required to demonstrate compliance with the regulations. Compliance must be shown for each combination of gross weight, center of gravity, altitude, temperature, airspeed, rotor RPM, etc.
Engineering tests are designed to investigate the overall capabilities and characteristics of the rotorcraft throughout its operational envelope. Testing will identify operating limitations, normal and emergency procedures, and performance information to be included in the FAA/AUTHORITY-approved portion of the flight manual. The testing must also provide a means of verifying that the rotorcraft’s actual performance, structural design parameters, propulsion components, and systems operations are consistent with all certification requirements.
(2) Section 21.35 requires, in part, that the applicant show compliance with the applicable certification requirements, including flight test, prior to official FAA/AUTHORITY Type Inspection Authorization (TIA) testing. Compliance in most cases requires systematic flight testing by the applicant. After the applicant has submitted sufficient data to the FAA/AUTHORITY showing that compliance has been met, the FAA/AUTHORITY will conduct any inspections, flight, or ground tests required to verify the applicant’s test results. FAA/AUTHORITY compliance may be partially determined from tests conducted by the applicant if the configuration (conformity) of the rotorcraft can be verified. Compliance may be based on the applicant’s engineering data, and a spot check or validation through FAA/AUTHORITY flight tests. The FAA/AUTHORITY testing should obtain validation at critical combinations of proposed flight variables if compliance cannot be inferred using engineering judgment from the combinations investigated.
(3) Performance tests include minimum operating speed (hover), takeoff and landing, climb, glide, height-velocity, and power available. Certain other performance tests, such as Category A, are conducted to meet specific requirements. Detailed performance test procedures and allowable extrapolation or simulation limits are contained in the respective paragraphs in this order.
(i) Hover tests are conducted to determine various combinations of altitude, temperature, and gross weight for both in-ground-effect (IGE) and, if required, Page B - 1
AC 29-2C, Chg 4 9/17/2009 out-of-ground effect (OGE) conditions. From these data the hover ceiling may be calculated.
(ii) Takeoff and landing tests are conducted to determine the total distance to takeoff and land at various combinations of altitude, temperature, and gross weight.
(iii) Climb tests establish the variations of rate-of-climb at the best rate-of-climb or published climb airspeed(s) at various combinations of altitude, temperature, and gross weight.
(iv) Height-velocity tests are conducted to determine the boundaries of the height versus airspeed envelope within which a safe landing can be accomplished following an engine failure.
(v) Power available tests are conducted to verify or reestablish the calculated installed specification engine performance model on which published performance is based.
(4) The purpose of rotorcraft stability and control tests is to verify that the rotorcraft possesses the minimum qualitative and quantitative flying qualities and handling characteristics required by the applicable regulations. In order to assess the handling qualities, standardized test procedures must be utilized and the results analyzed by accepted methods. Section 29.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 29.21(b) requires demonstration of controllability, stability, and trim. Combinations of §§ 29.21(a) and 29.21(b) may be used to show compliance to the operating envelope limits. Test methods and equipment are described in individual paragraphs of this advisory circular.
b. Procedures.
(1) Efforts should begin early in the certification program to provide advice and assistance to the applicant to insure coverage of all certification requirements. The applicant should develop a comprehensive test plan which includes the required instrumentation.
(2) The tests and findings specified in paragraph a(3) above are required of the applicant to show basic airworthiness and probable compliance with the minimum requirements specified in the applicable regulations. After these basic findings have been submitted and revi2ewed, a Type Inspection Authorization, or equivalent, can be issued. The FAA/AUTHORITY will develop a systematic plan to spotcheck and confirm that compliance with the regulations has been shown. The test plan will consider combinations of weight, center of gravity, RPM and cover the range of altitude and temperature for which certification is requested.
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AC 29.21A. § 29.21(Amendment 29-39) PROOF OF COMPLIANCE.
a. Explanation. Amendment 39 added § 29.83 which changes the requirements for determination of landing distance for Category B rotorcraft. This amendment requires landing distance to be determined with all engines operating within approved limits.
b. Procedures. The guidance material presented in paragraph AC 29.21 continues to apply.
AC 29.25. § 29.25 (Amendment 29-12) WEIGHT LIMITS.
a. Explanation.
(1) This section is definitive and specifies criteria for establishing maximum and minimum certificating weights. These weights may be based on those selected by the applicant, design requirements, or the limits for which compliance with all applicable flight requirements has been shown.
(2) Typical requirements that may establish the maximum and minimum weight limits include:
Maximum: Structural limits, performance requirements, stability, and controllability requirements.
Minimum: Autorotative rotor RPM, stability, and controllability requirements.
(3) Jettisonable External Cargo.
(i) Paragraph (c) was added by Amendment 29-12 to provide, in the certification standards, a basis for approving an increase in gross weight (exceed standard limits) that would be an external jettisonable load. The attachment device standards were moved from Part 133 (Amendment 133-5) to Parts 27 and 29.
Section 29.865, “External load attaching means,“ now contains the standards, including design features, for the attaching devices. Cargo hoists and hooks were envisioned.
Prior to these amendments, type design approvals were made under Part 133 and the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 whenever the standard limits were exceeded.
(ii) In the preamble of Amendment 29-12 (Proposal 2-99, 41 FR 55454, December 20, 1976) the agency stated, in part, that “...§ 29.25(c) is intended to provide only a total weight standard for approving the rotorcraft structure (and propulsion systems) for operation under Part 133.” As indicated in § 29.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, transmissions, engines, etc., are subject to evaluation under § 29.571 for external cargo approval whenever the “standard” structural limitations are exceeded (Review Case Nos. 37 and 55).
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AC 29-2C, Chg 4 9/17/2009
(iii) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering approval. This Part 133 standard is also applicable for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM supplement for external load operations to prepare a rotorcraft load combination flight manual required by § 133.47.
b. Procedures.
(1) It may not be possible to demonstrate quantitatively all the flight requirements at the minimum weight because of test instrumentation requirements. The test team must be assured that the rotorcraft complies with the applicable requirements at the lowest permissible flying weight. This evaluation may be done qualitatively, with the test instrumentation removed, and with minimum crewmembers if no critical areas exist or are anticipated. Additionally, reasonable extrapolation may be warranted. Howev er, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted.
(2) Whenever a gross weight increase (§ 29.25(c)) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo shapes and weights (density) is satisfactory. All possible combinations of weights and shapes are not evaluated. The representative configurations may be noted in the RFM or RFM supplement for the operator’s information. Sections 133.41 and 133.47 must be satisfied by the individual operator for the particular case at hand. The approved RFM or RFM supplement should provide the necessary limitations and any other information about the representative cargo configurations evaluated. Section 133.41 also permits the operator to obtain approval of additional and unique cargo configurations provided the approved limitations are observed. Paragraph AC 29.1581 concerns the RFM and its contents.
(3) See paragraph AC 29.571, § 29.571, for fatigue substantiation and external cargo considerations.
(4) Refer to AC 133-1A, Rotorcraft External-Load Operations in Accordance with FAR Part 133, October 16, 1979, for further information on airworthiness and flight manual policy.
AC 29.25A. § 29.25 (Amendment 29-51) WEIGHT LIMITS.
a. Explanation.
(1) Amendment 29-51 added a new paragraph (a)(4) that requires that the operating envelope for the controllability demonstrated under § 29.143(c) be included in the limitations section of the Rotorcraft Flight Manual (RFM). The change allows, in addition to the 17-knot controllability requirements, the applicant to provide additional Page B - 4
controllability information within an applicant selected limited azimuth range if the rotorcraft is certified with nine or less passenger seats. This effectively allows increased weights within this limited range. Amendments 29- 21 and 29-24 allowed for this relief and subsequent regulatory policy recognized these limitations as they are now required.
In no case should those limits be established at an altitude that is not operationally suitable. In the past, the minimum operationally suitable altitude for takeoff and landing has been established as 3,000 feet density altitude.
(2) The explanation regarding the relief for presentation of hover controllability limits in AC 29.143.a.(2)(ii) (Amendment 29-24) is superseded by this change.
b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect.
AC 29.27. § 29.27 (Amendment 29-3) CENTER OF GRAVITY LIMITS.
a. Explanation.
(1) This regulation is definitive and requires that the center of gravity limits be defined. Proof of compliance with all applicable flight requirements is required within the range of established CG’s. Along with the longitudinal CG limits, the lateral CG limits should either be established or determined to be not critical.
(2) Ballast is usually carried during the flight test program to investigate the approved gross weight/center of gravity limits. Lead is the most commonly used form of ballast during rotorcraft flight testing although other types of ballast, such as water, may serve just as well. Water may have the added benefit of being jettisonable during critical flight test conditions. Care must be taken regarding the location of ballast. The strength of the supporting structures should be adequate to support such ballast during the flight loads that may be imposed during a particular test and for the ultimate inertia forces of § 29.561(b)(3). Of critical importance is the method of securing the ballast to the desired locations. To avoid any undesired in- flight movements of the ballast, a positive method of constraint is mandatory. The flight test crews should also visually verify the amount, location, and integrity of the ballast. The effects of mass moment of inertia on the flight characteristics due to the ballast locations should also be considered. The mass moment of inertia of the test rotorcraft should, to the extent possible, be the same as that expected in normal, approved loadings, especially during tests involving dynamic inputs.
b. Procedures.
(1) Center of gravity locations and limits are of prime importance to rotorcraft stability and safety of flight. The primary concern is establishment of the longitudinal center of gravity limits. Lateral center of gravity limits with respect to longitudinal center of gravity limits are also important. The design of the rotorcraft is usually such that approximate lateral symmetry exists. This lateral symmetry can be upset by lateral Page B – 5
AC 29-2C, Chg 4 5/1/2014 loadings resulting in the necessity to establish lateral center of gravity limits. There are two characteristics which may be seriously affected by loading outside the established center of gravity limits; these are stability and control. The established center of gravity limits must be such that as fuel is consumed, it is possible for the rotorcraft to remain within the established limits by acceptable loading and/or operating instructions.
(2) Structural limits may restrict the maximum forward longitudinal center of gravity limits. However, in most cases it is the maximum value established wherein adequate low speed control power exists to meet such requirements as § 29.143(c).
Likewise, the maximum aft center of gravity limit may be a “structural limit,” but it usually is determined during flight test after the rotorcraft’s handling qualities tests have been conducted. Additional items which may influence the maximum aft center of gravity limits may be malfunctions of automatic stabilization equipment, excessive rotorcraft attitudes during critical phases of flight, or adequate control power to compensate for an engine failure.
(3) Lateral center of gravity limits have become more critical because of the ever increasing utilization of the rotorcraft for such things as unusual and unsymmetric lateral loads, both internal and external. Maximum allowable lateral center of gravity limits have also influenced the results of the unusable fuel determination.
(4) Summarizing, it is of prime importance that longitudinal and lateral center of gravity limits be determined so that unsafe conditions do not exist within the approved altitude, airspeed, ambient temperature, gross weight, and rotor RPM ranges. All relevant malfunctions must be considered.
AC 29.29. § 29.29 (Amendment 29-15) EMPTY WEIGHT AND CORRESPONDING CENTER OF GRAVITY.
a. Explanation. The empty weight of the rotorcraft consists of the airframe, engines, and all items of operating equipment that have fixed locations and are permanently installed (including both required and optional equipment) in the rotorcraft. It includes fixed ballast, unusable fuel, other unusable fluids, and full operating fluids. “Full operating fluids” such as oils used in an engine, auxiliary power unit, main and auxiliary gearboxes, and hydraulic systems are considered “closed fluid systems” typically filled to a “full mark” indicator level. Fluids necessary for the operation of non- permanently installed equipment (i.e., carry-on equipment) are not considered part of the empty weight.
(1) A ballast is fixed when made a permanent part of the rotorcraft as a means of controlling the certificated empty weight center of gravity (CG).
(2) Installed equipment is any FAA-approved equipment attached to the rotorcraft with hardware and, as a result, becomes an integral part of the rotorcraft. The installation or removal of such equipment must be recorded in the aircraft equipment list. Compliance with paragraph (b) of § 29.29 is accomplished by the use of an Page B – 6
equipment list specifying the installed equipment at the time of weighing and the weight moment arm of the equipment.
b. Procedures.
(1) Determination of the empty weight and corresponding center of gravity is primarily the responsibility of the manufacturer and is normally made on a production rotorcraft rather than a prototype. If the manufacturer has been issued a production certificate and wishes to avoid weighing each production rotorcraft, the manufacturer may make a detailed proposal defining the procedure it would use to establish an empty weight and CG. When the proposal is approved, the manufacturer will weigh the first five to ten production rotorcraft and show that the rotorcraft will be within ±1 percent on empty weight and ±0.2 inches on CG. After this procedure is established, the empty weight and CG may be computed except that at regular intervals a rotorcraft will be weighed to ensure the tolerances are still being maintained (e.g., one in ten rotorcraft).
(2) For prototype and modified rotorcraft, it is only necessary to establish a known basic weight and CG position (by weighing) from which the extremes of weight and CG travel required by the test program may be calculated. See the current version of FAA-H-8083-1 (Pilots Weight and Balance Handbook) for a sample weight and balance procedure.
(3) The weight and balance should be recalculated if a modification (or series of modifications) to the rotorcraft results in a significant change to the empty weight.
Additionally, this change in empty weight should be reflected with the weight and balance information contained in the rotorcraft flight manual or rotorcraft flight manual supplement.
c. Ballast Loading and Type.
(1) Ballast loading of the rotorcraft can be accomplished in any manner to achieve a specific CG location. It is acceptable for such ballast to be mounted outside the physical confines of the rotorcraft if the flight test objectives are not affected by this arrangement. In flight test work, loading problems will occasionally be encountered in which it will be difficult to obtain the desired CG limits. Such cases may require loading in engine compartments or other places not designed for load carrying. When this condition is necessary, care should be taken to ensure that local structural stresses are not exceeded or that the rotorcraft flight characteristics are not changed due to increased moments of inertia by attaching the ballast to extreme CG locations that may not be designed for the added weight.
(2) The two types of ballasts that may be used in loading are solid or liquid. The solids are usually high density materials such as lead, while the liquid usually used is water. In critical tests, the ballast may be loaded in a manner so that disposal in flight can be accomplished. In any case, the load should be securely attached in its loaded position so shifting or interference with safety of flight will not result.
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AC 29-2C, Chg 4 5/1/2014 AC 29.31. § 29.31 REMOVABLE BALLAST.
a. Explanation. This regulation provides the option of using removable ballast for operational flights to obtain center of gravity locations that are in compliance with the flight requirement of this Part. Fixed ballast used for flight operations after type certification must be documented in the type design data. Removable ballast is used primarily on small rotorcraft to control the CG with different passenger loadings although this regulation does permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations.
See paragraph AC 29.873 for information on ballast provisions.
b. Procedures. None.
AC 29.33. § 29.33 (Amendment 29-15) MAIN ROTOR SPEED AND PITCH LIMITS.
a. Explanation.
(1) General. This rule requires the establishment of power -on and power-off main rotor speed limits and the requirements for low rotor speed warning.
(2) Power-On. The power -on limits should be sufficient to maintain the rotor speed within these limits during any appropriate maneuver expected to be encountered in normal operations throughout the flight envelope for which certification is requested.
A power-on range of approximately 3 percent has in the past been the minimum range required due to engine governor and engine operating characteristics. With the introduction of advanced engines and electronic engine controls, there may not be a need for a range, but one fixed value may suffice. Transient power -on values may also be acceptable provided they are substantiated.
(3) Power-Off. The power -off rotor speed limits should be sufficient to encompass the rotor speeds encountered during normal autorotative maneuvers except for final landing phase (touchdown) for which rotor RPM may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. The limits should also be sufficient to cover the ranges of airspeed, weight, and altitudes for which certification is requested. It is not the intent of the rule to require the minimum and maximum limit values in conjunction with extremes such as maximum/minimum weights and/or high altitude. The minimum and maximum rotor speed requirements should be thoroughly evaluated at normal operation environment; i.e., at altitudes between approximately sea level and 10,000 feet, temperatures not at extremes, and weights as necessary for other tests and as required to readily establish the limit rotor speeds.
Spot checks of the autorotative requirements should be made at the extremes of the flight envelope and environmental conditions during normal tests at those conditions.
Under conditions where high autorotative rotor speeds may be encountered, it is acceptable for the pilot to adjust the controls to prevent overspeeding of the rotor. At light weight combined with low altitudes and extreme cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within Page B – 8
limits. If this oc curs, the manufacturer may elect to adjust the low pitch stops as a maintenance procedure at extreme ambient conditions provided the flight and maintenance manuals clearly present the rigging requirements and procedures. There must be sufficient “overlap” of ambient conditions between configurations such that rerigging is not required whenever ambient temperature and surface elevation change slightly. Any down rigging of the low pitch stop must continue to ensure adequate clearance between controls and other rotorcraft structure and should be evaluated during flight test. Both the power -on and power-off limits may also be established by encountering critical flapping limits in some approved flight conditions such as high airspeed or sideward flight.
(4) Additional RPM Ranges. Some applicants have elected to certify their aircraft with additional RPM ranges in an attempt to realize additional performance during certain flight conditions or maneuvers such as Category A OEI continued and rejected takeoffs and balked landings. Such additional RPM ranges have been found acceptable as long as all pertinent FAR requirements are fully substantiated for operation in that range. The substantiation should include drive system endurance and flight test verification of performance and flight characteristics during applicable maneuvers, in the additional RPM range. The FAA/AUTHORITY does not define additional RPM ranges as transient since all applicable requirements must be satisfied for approval of that range.
(5) Low Speed Warning. If it is possible under expected operating conditions for the rotor speed to fall below the minimum approved values, the requirement exists for a low rotor speed warning. This warning is required on all single- engine rotorcraft and on multiengine rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although today’s multiengine rotorcraft do not require a low rotor speed warning according to the rule, essentially all have warning systems installed. If the minimum power -on and power-off rotor speed limits are different, the warning signal should be at the higher speed, normally the power-on minimum rotor speed. One rotorcraft has a warning system cutout if the collective is full down, and others have other warnings on the engine speed to indicate engine failure.
All of these related warning systems must be evaluated with emphasis on ensuring adequate rotor speed.
b. Determination and Testing. Refer to paragraph AC 29.1509 (§ 29.1509) for additional information on rotor limits determination and testing.
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AC 29-2C, Chg 4 5/1/2014
SUBPART B - FLIGHT
PERFORMANCE AC 29.45. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation.
(1) Changes to various part 29 sections, which did not include amending § 29.45, added new and revised airworthiness standards for the performance of transport category rotorcraft and renumbered several sections within the performance section of Subpart B. The performance section of this guidance material has been organized for easy use with rotorcraft certificated before or after this amendment. To achieve this, some of the guidance material has been duplicated under different paragraph numbers.
A statement at the beginning of each of these paragraphs indicates where other pertinent information can be found.
(2) Section 29.45 lists the rules and standards under which the performance requirements are to be met. This guidance provides general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables that affect the flight test results that must be taken into account. Some of these variables are wind, temperature, altitude, humidity, rotorcraft weight, power, rotor RPM, center of gravity, etc. A thorough knowledge of the testing procedures and data reduction methods is essential and good engineering judgment must be used to determine applicable test conditions. The test results should be analyzed and expanded by an approved methodology. The guidance within this section is considered an approved methodology.
(3) Performance should be based on approved engine power as determined in paragraph b.(4) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits. Any sustained operation in these transient limit areas usually require some form of special maintenance. However, for such demonstrations as rejected and continued OEI category A takeoffs and HV determination, low rotor speeds have been authorized based upon additional structural and drive system substantiation (see section 29.33 of this AC).
(4) Where variations in the parameter on which a tolerance is allowed will have an appreciable effect on the test, the results should be corrected to the standard value of the parameter; otherwise, no correction is necessary.
(5) As defined in 14 CFR § 1.1, the 30-second and 2-minute OEI power ratings are based on up to three periods of use during a single flight. The purpose of the three applications is: (i) to initially recover from an engine failure, (ii) to conduct OEI missed approach, and (iii) to conduct the final OEI landing. Rotorcraft performance based on Page B – 10