44. § 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 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 52 Par 44
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.
45. § 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 in the rotorcraft. 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 certificated empty weight CG.
(2) Compliance with Paragraph (b) of § 29.29 is accomplished by the use of an equipment list which defines 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 a manufacturing inspector and is normally made on a production rotorcraft rather than a prototype. If a manufacturer wishes to avoid weighing each production rotorcraft and has been issued a production certificate, the manufacturer may make a detailed proposal defining the procedures used 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 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 91-23A, Pilots Par 44 53
Weight and Balance Handbook, June 9, 1977, 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.
(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) Two types of ballast that may be used in loading 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. § 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 385 of this advisory circular for information on ballast provisions.
b. Procedures. None.
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7/30/97 АС 29-2B
47. § 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 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 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.
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(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 721 (§ 29.1509) for additional information on rotor limits determination and testing.
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48.-54. RESERVED.
Intentionally Left Blank Par 48 57 (thru 63)
SECTION 3. PERFORMANCE
55. § 29.45 (Amendment 29-24) GENERAL.
a. Explanation.
(1) Amendment 29-39 adopts new and revised airworthiness standards for the performance of transport category rotorcraft. As part of this change, several sections within the performance section of Subpart B were renumbered. The performance section of this guidance material has been organized for easy use with rotorcraft certified 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 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. The test results should be analyzed and expanded by approved methodology within the guidelines of this paragraph. 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.
(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 OLI Category A takeoffs and height-velocity (HV) determination, low rotor speeds have been authorized based upon additional structural and drive train substantiation (see Paragraph 47).
(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.
b. Procedures.
(1) Winds For Testing.
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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 distance tests than for hover performance. Likewise higher winds can sometimes be tolerated during hover performance testing on large, heavy rotorcraft with high rotor downwash velocities than for smaller rotorcraft with 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.
(ii) Past experience has shown that a steady wind of 0 to 10 knots will result in acceptable takeoff and landing performance if distances are corrected for the winds measured during these tests. This is not the case for vertical takeoffs and landings. To obtain consistent and repeatable vertical performance data, the same general wind criteria used to obtain hover performance; i.e., up to 3 knots, should be adhered to for vertical performance determination. In actuality, a rotorcraft may exhibit reduced IGE hover performance in winds from 3 to 15 knots due to partial immersion of the main rotor in its own vortex. Since the height-speed envelope determination is affected by wind just as vertical takeoff and landing performance are, the same allowable winds for testing should be adhered to for HV testing; i.e., O to 3 knots.
(Ili) 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 are given for general guidance only:
(A) Hover performance - 0 to 3 knots.
(B) Conventional takeoff and landing - 0 to 10 (data to be corrected) (C) Vertical takeoff and landing - 0 to 3 knots (D) Height-velocity - 0 to 3 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. The wind effects on required runway length for takeoff and landing distances may be shown in the flight manual.
(V) Full wind credit may be given for conventional takeoff and landing field lengths. This credit should not be more than the nominal wind component along the takeoff or landing path opposite to the direction of flight.
(2) Altitude Effects. Using FAA/AUTHORITY-approved methodology, hover, takeoff, and landing, performance may be extrapolated and/or interpolated from test data up to a maximum of 14,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be Par 55 65
extrapolated more 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 § 29.1527 and delineates the maximum altitude up to which operation is allowed.
This altitude normally constitutes the maximum cruise or enroute altitude.
(C) Maximum weight, altitude and temperature for takeoff and landing constitutes a limitation. 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 §§ 29.51, 29.53, 29.59, 29.63, 29.73, 29.1583 and 29.1587.
(ii) Procedures.
(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required:
(1) Takeoff (§§ 29.51-29.63)
(2) Climb (§§ 29.64-29.67)
(3) Performance at minimum operating speed (§ 29.49)
(4) Landing (§ 29.75)
(5) Limiting height-speed envelope (§ 29.87)
(6) IGE controllability (§ 29.143c) 66 Par 55
(L) Cooling (§§ 29.1041-29.1045)
(8) Engine operating characteristics (§ 29.939) 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.
(C) Prior to Amendment 29-21, H-V information was an operating limitation. With the adoption of Amendment 29-21, the H-V curve is performance information for Category B rotorcraft with nine or less passenger seats but remains a limitation for Category A rotorcraft and Category B rotorcraft with 10 or more passenger seats.
(D) Prior to Amendment 29-24, IGE controllability was required in 17 knots of wind to the maximum takeoff and landing conditions. With the adoption of Amendment 29-24, if IGE or OGE hover performance is presented for a Category B rotorcraft to an altitude in excess of that for which IGE controllability at 17 knots is presented, the maximum safe wind demonstrated for hover operations must be presented in the RFM. The amendment did not change the requirement for Category A rotorcraft.
(E) The requirements for data collection and presentation in the RFM vary depending upon the certification basis of the rotorcraft. These requirements are presented by regulation and amendment in Figures 55-1 and 55-2.
(F) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in Paragraph b(2) above and not above the lowest limiting altitude resulting from the requirements listed in subparagraph A of this paragraph.
(4) Temperature Effects.
Background.
(A) The regulations prohibit any unsafe design feature throughout the range of environmental conditions for which certification is requested. The regulations also require that the performance and handling qualities be determined over the approved range of atmospheric variables selected by the applicant.
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(B) Substantiation of temperature effects on performance and handling characteristics is required throughout the approved temperature range. In the past, approved analyses were frequently accepted for determining the extreme temperature effects on performance and flight characteristics. With the introduction of newer, higher performance rotorcraft, advanced rotor blade designs, higher airspeeds, and blade mach numbers, the previous methods have proven to be insufficient.
Therefore, the performance and flight characteristics should be validated at extreme temperatures; however, analysis may be permitted if a suitable methodology is demonstrated.
(C) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected, sometimes significantly, in both the performance and flying qualities areas. Hot temperature conditions although not shown to be as critical should be given consideration.
(D) Additionally, design deficiencies surfaced when the rotorcraft were exposed to temperature extremes and some of these difficulties were severe enough to require the redesign of equipment and/or materials. Therefore, to satisfy § 29.1309(a), the applicant needs to substantiate the total rotorcraft at the extreme temperatures for which certification is requested.
(її) Procedures.
(A) The FAA/AUTHORITY is responsible for verifying the applicant's predictions of performance and handling characteristics at the temperature extremes for which certification is requested. A limited flight verification, if necessary, could include spot checks of hover and climb performance, IGE controllability, roughness determination, simulated power failure, static stability, height-velocity, VENo evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.
(B) Extrapolation of test data should only be allowed if the applicant's predicted or calculated data is verified by actual test but in any case extreme caution should be used for extrapolations that are -10°C below or +20° C above those values tested.
(5) Weight Effects. Test weights should be maintained within +3 percent and -1 percent of the target weight for each data point. Weight may be extrapolated only along an established W/o line within the allowable altitude extrapolation range.
(6) Engine Power - Turboshaft Engine:
(1) Background.
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(A) The purpose of rotorcraft performance flight testing is to obtain accurate quantitative flight test performance data to provide flight manual information.
(B) Flight tests are designed to investigate the overall performance capabilities of the rotorcraft throughout its operating envelope. This testing furnishes information to be included in the flight manual and provides a means of validating the predicted performance of the rotorcraft with a minimum installed specification engine.
(C) The horsepower used to complete the flight manual performance must be based on horsepower values no greater than that available from the minimum uninstalled specification engine after it is corrected for installation losses. A minimum uninstalled specification engine is one that, on a test stand under conditions specified by the engine manufacturer, will produce the certificated horsepower values at specification temperatures and/or speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specified horsepower at a particular engine temperature or speed rating with higher allowable limits. The limit is the maximum value the installed engine is allowed in order to develop the specification horsepower. Prior to installation of each engine in a rotorcraft, the performance is measured by the engine manufacturer. This is done by making a static test run in a test cell and referring the results to standard day, sea level conditions. The performance parameters obtained are presented as uninstalled engine characteristics on a test log sheet. This is commonly referred to as a "final run sheet." Figure 55-3 compares a typical engine to one the manufacturer has certified as a minimum uninstalled certified engine.
(D) After engine certification, the engine manufacturer is responsible to ascertain that each engine delivered will produce, as a minimum, the certified horsepower values without exceeding specification operating values; therefore, a "final run sheet" is created for every engine produced. Additionally, if needed, arrangements can usually be made with the engine manufacturer to obtain a torque system calibration for individual engines. This will further optimize the accuracy of the engines used in the flight test program. The engine manufacturer will also provide predicted uninstalled power available for the various power ratings. This information may be derived from an engine computer "card deck" and from charts and tables in the engine detail installation manual. These data also provide engine performance for the range of altitudes and temperatures approved for the engine and include methods for correcting this performance for installation effects. The parameters contained in a typical "card deck" are plotted for one engine rating in Figure 55-7.
(E) Several power losses may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators. Additional flight manual performance considerations are the torque indicating system accuracy and torque needle split. The predicted uninstalled power available engine characteristics cannot be assumed to be the actual power available after the engine is installed in the Par 55 69