(її) Takeoff and landing tests are conducted to determine that a takeoff or landing can be safely executed without requiring exceptional piloting skill or favorable conditions at any approved combination of altitude, temperature, and gross weight.
(ili) For rotorcraft other than helicopters, 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. For helicopter, climb tests are conducted as required to determine the best rate-of-climb speed, Vy.
(iv) Height-velocity tests are conducted to determine the boundaries of the height versus airspeed envelope from which a safe landing can be accomplished following an engine failure.
(V) Power available tests are conducted to verify 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 27.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 27.21(b) requires demonstration of controllability, stability, and trim. Combinations of § 27.21(a) and (b) may be used to show compliance with 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 ensure 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 42a(3) 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 reviewed, 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, and RPM and cover the range of altitude and temperature for which certification is requested.
43. § 27.25 (Amendment 27-14) WEIGHT LIMITS.
a. Explanation.
Par 42 44
(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:
(i) Maximum: Structural limits, performance requirements, stability, and controllability requirements.
(II) Minimum: Autorotative rotor RPM, stability, and controllability requirements.
(3) Jettisonable External Cargo.
(i) Section 27.25(c) was added by Amendment 27-11 to provide a basis for approving an increased gross weight that would be an external jettisonable load.
Section 27.865, "External load attaching means, " includes hoist and hook design features for the load attaching devices that were added to Part 27 but removed from § 133.43. Part 133, "Rotorcraft External-Load Operations," was also amended (Amendment 133-5) concurrently to complement the changes to Parts 27 and 29.
(Ш) Approvals under the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 were no longer necessary. These review cases concerned the policy/standards for external cargo configurations using a cargo hook whenever the standard limitations were exceeded. If the standard limitations were not exceeded, external cargo hooks and hoists and external cargo configuration approvals could be made under Part 133, Subpart D, prior to Amendment 133-5.
In the preamble of Amendment 27-11 (Proposal 2-99, 41 FR 55454;
December 20, 1976) the agency stated that"...§ 27.25(c) and § 29.25(c) are intended to provide only a total weight standard for approving the rotorcraft structure for operation under Part 133." The policy in Review Case No. 55 also indicates the powerplant or propulsion system is also subject to evaluation for the increased weight.
As indicated in § 27.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, etc., are only subject to fatigue evaluation under § 27.571 whenever the standard structural limitations are exceeded (Review Case No. 55).
(iv) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering approval. Section 133.41 is also appropriate for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM Par 43 45
supplement to prepare his own 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 ensure 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 is permitted. However, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted.
(2) Whenever a gross weight increase under § 27.25(c) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo weights and/or shapes (or 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 approved limitations are observed. Paragraph 762 of this AC concerns the RFM and its contents.
(3) See AC 29-2, Certification of Transport Category Rotorcraft Paragraph 230, concerning § 29.571, for fatigue substantiation and external cargo considerations that apply to § 27.571 as well.
(4) Refer to AC 133-1A, Rotorcraft External-Load Operations in Accordance with FAR Part 133, for further information on airworthiness and flight manual policy for operators.
44. § 27.27 (Amendment 27-2) 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 Par 43 46
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 in 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 numerous probable lateral loadings possibly resulting in the necessity to establish lateral center of gravity limits. Stability and control characteristics may be seriously affected by loading outside the established center of gravity limits. The established center of gravity limits must be 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 § 27.143(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. Flight tests may reduce the "structural limit" CG envelope, but flight tests alone should not be used to expand the "structural limit." 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) In summary, 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 Par 44 47
altitude, airspeed, ambient temperature, gross weight, and rotor RPM ranges. All relevant malfunctions must be considered.
45. § 27.29 (Amendment 27-14) 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 in the aircraft. It includes fixed ballast, unusable fuel, and full operating fluids except water intended for injection in the engines.
(1) Fixed ballast refers to ballast that is made a permanent part of the rotorcraft as a means of controlling the empty weight CG.
(2) Compliance with Paragraph (b) of § 27.29 is accomplished by the use of an equipment list which defines the installed equipment at the time of weighing and the weight arm and moment of the equipment.
b. Procedures.
(1) Determination of the empty weight and corresponding center of gravity is primarily the responsibility of the manufacturing inspector. This determination is normally made on the production rotorcraft rather than the prototype. If the manufacturer wishes to avoid the necessity of weighing each production rotorcraft and he has been issued a production certificate, he may make a detailed proposal defining the procedures he will use to establish an empty weight and CG When his proposal is approved, he will weigh the first five to ten production rotorcraft and show that the rotorcraft will be within 11 percent on empty weight and t0.2 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 AC 91-23 (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 (RFM) or Rotorcraft Flight Manual Supplement (RFMS).
c. Ballast Loading and Type.
Par 44 48
(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 which may not be designed for the added weight.
(2) There are basically two types of ballast that may be used in loading. They are solids 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.
46. § 27.31 REMOVABLE BALLAST.
a. Explanation. This regulation provides the option of using removable ballast to obtain desired center of gravity locations to determine 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 not permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations.
b. Procedures. None
47. § 27.33 (Amendment 27-14) MAIN ROTOR SPEED AND PITCH LIMITS.
a. Explanation.
(1) General. This section 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.
In the past a minimum power-on range of approximately 3 percent has been 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.
One fixed value may suffice. If substantiated, transient power-on values may also be acceptable.
Par 45 49
(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 operating environment; l.e., at altitudes from approximately sea level to 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 extremely cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within limits. If this occurs, 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 downrigging 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 tight.
(4) 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 not required by the rule, essentially all of today's multiengine rotorcraft have a low rotor speed warning system 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 type of rotorcraft has a warning system cutout if the collective is full down, and other types 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 721 (§ 27.1509) for additional information on determining and testing rotor limits.
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48.-56. RESERVED.
Intentionally Left Blank Par 48 51 (thru 59)
AC 27-1A 7/30/97/
57. RESERVED.
Intentionally Left Blank Par 57 60
7130/97 AC 27-1A SECTION 3. PERFORMANCE
58. § 27.45 (Amendment 27-21) GENERAL
a. Explanation.
(1) Section 45 of Part 27 lists some of the rules and standards under which the performance requirements are to be met. This paragraph will provide general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables which 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 acceptable test conditions. The test results should be analyzed and expanded by approved methodology within the guidelines of this paragraph.
(2) Performance should be based on approved engine power as determined in Paragraph b(5) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits and any sustained operation in these transient limit areas usually requires some form of special maintenance. However, for such demonstrations as landing procedure demonstration and height-velocity (HV) determination, low rotor speeds (within approved limits) have been authorized. Such transients, if authorized, must be flight evaluated.
(3) 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.
b. Procedures.
(1) Winds for Testing.
Allowable wind conditions will vary with the type of test and will also be different for different types and gross weight rotorcraft. For example, higher winds can usually be tolerated for takeoff and landing tests than for hover performance.
Higher winds can sometimes be tolerated during hover performance testing on rotorcraft with high rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and/or accounted for, hover performance testing should be conducted in winds of 3 knots or less.
(i¡) In-ground-effect controllability and maneuverability testing should be conducted in surface winds of less than 5 knots, or when higher steady wind conditions exist, with a maximum gust spread of 5 knots.
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АС 27-1A 7/30/97 As can be seen from the foregoing, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions is given for general guidance only:
(A) Hover performance - 0 to 3 knots.
(B) Height-velocity - 0 to 3 knots.
(C) IGE controllability and maneuverability - 0 to 5 knots.
(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests.
(2) Altitude Effects. Using FAA/AUTHORITY-approved methodology, hover performance may be extrapolated and/or interpolated from test data up to a maximum of +4,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be extrapolated higher than approximately 2,000 feet density altitude from the test altitude. Cruise stability/controllability tests should be evaluated at least at two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet. As in all testing, extrapolation and/or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions.
(3) Altitude Limitations.
Explanation.
(A) Two altitudes are normally presented in the RFM to define the operating envelope of a rotorcraft;
- Maximum operating altitude, and
- Maximum takeoff and landing altitude.
(B) Maximum operating altitude is an operating limitation required by § 27.1527 and delineates the maximum altitude to which operation is allowed. This altitude normally constitutes the maximum cruise or en route altitude.
(C) Maximum takeoff and landing altitude is the hover in-ground-effect (IGE) ceiling for a rotorcraft as described in § 27.73. The hover ceiling and any information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certified to CAR 6, Amendment 6-7 or any amendment of FAR 27, a hover ceiling may not be presented above the altitude at which H-V and IGE controllability tests were conducted plus allowable extrapolation, Par 58 62
unless that extrapolated altitude is at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, then that altitude is the maximum takeoff and landing altitude of the rotorcraft. The maximum takeoff and landing altitude may be coincident with, but never above the maximum operating altitude limitation. Takeoff and landing and hover ceiling data and presentation requirements are presented in §$ 27.51, 27.73 and 27.1587.
(її) Procedures.
(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required:
(1) Takeoff (§27.51)
(2) Climb (§§ 27.65 and 27.67)
(3) Performance at minimum operating speed (§ 27.73)
(4) Landing (§ 27.75)
(2) Limiting height-speed envelope (§ 27.79)
(6) GE controllability (§ 27.143c) (Z) Cooling (§§ 27.1041, 27.1043 and 27.1045)
(8) Engine operating characteristics (§ 27.939) Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this AC.
(B) As detailed in subparagraph b(2) above, the maximum allowable extrapolation of H-V, IGE controllability and engine operating characteristics is +2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site, or for CAR 6, Amendment 6-7 and subsequent, unless test results were demonstrated to at least 7,000 feet.
(C) If IGE controllability is demonstrated to at least 17 knots of wind at 7,000 feet, hover capability above 7,000 feet may be presented provided that the maximum demonstrated safe wind for takeoff and landing above 7,000 feet is specified in the RFM.
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