(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.
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(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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АС 27-1А 7/30/97 (D) 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 58-1 and 58-2.
(E) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in Paragraph b(2) and not above the lowest limiting altitude resulting from the requirements of subparagraph A of this paragraph.
(4) Temperature Effects.
Background.
(A) 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 higher blade tip 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.
(B) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected by cold temperature in both the performance and flying qualities areas. Hot temperature conditions, although not shown to be as critical for flying qualities, should be given consideration.
(C) 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 § 27.1309(a), the applicant needs to substantiate the total rotorcraft throughout the foreseeable range of operating temperatures.
(її) Procedures.
(A) The FAA/AUTHORITY is responsible for verifying the effects of temperature on performance and handling characteristics. A limited flight verification, if necessary, could include spot checks of hover performance, IGE controllability, vibration, simulated power failure, static stability, height-velocity, VNE/No 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.
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7/30/97/ АС 27-1A
(5) Engine Power - Turboshaft Engine.
(i) Background.
(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 power used to complete the flight manual performance must be based on power 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 power at specification temperatures and/or speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specitied power 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 power. 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 58-3 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 certitied power 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 58-4.
(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 Par 58 65
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 rotorcraft because this procedure would neglect the installation power losses. It is necessary to know the installation losses in order to determine the flight manual performance. Installation losses are reflected reductions in available power resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet and/or exhaust design. The rotorcraft manufacturer conducts tests to confirm the installed specification engine power available on which published performance is based. The specific methods used vary widely between manufacturers but usually include some combination of ground and flight tests. Figure 58-5 is a typical installed power available chart for one set of conditions.
(F) The installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion may occur and there are changes in airflow routing. The rotorcraft manufacturer may elect to determine the installation losses for different flight conditions to take any airspeed advantages. This is acceptable if, for example, the hover performance is based on the actual power available from an installed minimum specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual power available from an installed minimum specification engine at the published climb airspeed. This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.
(її) Procedures.
(A) The installed minimum specification engine power output has been predicted and calculated for various flight conditions. It is imperative that the predicted values be verified by actual flight test. The flight test involves obtaining engine performance measurements at various power settings, altitudes, and ambient temperatures. The data should be obtained at the actual flight condition for which the performance is to be presented (i.e., hover, climb, or cruise).
(B) Following a power increase, engine temperature and/or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM and/or temperature are held constant. This is a characteristic typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities. An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data.
(1) To determine the takeoff and 2 ½-minute values, first stabilize the engine at a low power setting. After stabilization, rapidly increase the power Par 58 66
demand to takeoff and/or 2 ½-minute power levels. Record the engine parameters as soon as the specification torque, temperature, or speed is attained. Care must be taken not to exceed a limit. These readings should be obtained approximately 15 seconds after power is initially applied.
(2) To determine the 30-minute and/or maximum continuous power values, approximately 2 to 3 minutes of stabilization time after power is increased is generally used, but up to 5 minutes stabilization time is allowed. The reason for the different procedures is when a pilot requires takeoff or 2 ½-minute power values he is in a critical flight condition and does not have the luxury of waiting for the engine(s) to produce rated power. Stabilization time is allowed for the maximum continuous and 30-minute ratings because these values are not associated with flight conditions for which power is needed immediately.
(C) The in-flight measurements recorded with the engine(s) on the flight test rotorcraft must be corrected downward if the test engine is above minimum specification and corrected upward for a test engine that is below minimum specification. This correction is necessary to verify that a minimum specification engine installed on a production rotorcraft is capable of producing the power values used to compute the flight manual performance without exceeding any engine limit. In addition, if the production rotorcraft's power measurement devices have significant (greater than 3 percent) power error, this error must be accounted for in a conservative manner.
(D) On multiengine rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengine performance should be based on the total power available after considering the different installation losses and with minimum specification engines installed.
One-engine-inoperative performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, engine driven accessories, etc., must be accounted for accordingly.
(E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical No speed to a constant corrected speed at cold temperatures. Others may limit power to a fuel flow value which would be encountered only at certain ambients.
Others may limit by torque limiting devices. Therefore, power available data should be obtained at various ambients to verify that all limiting devices are functioning properly and have not been affected by the installation.
(F) Through use, turbine engine power capabilities decrease with time. This is called engine deterioration. Deterioration is largely a function of the particular engine design, the manner, and the environment in which the engine is operated. There is a need, therefore, to provide a method which can be used in service to periodically determine the level of engine deterioration. A power assurance curve is Par 58 67
AC 27-1A 7130/97 usually provided to allow the flightcrew to know the power producing capabilities of any engine. A power assurance check is a check of the engine(s) which will determine that the engine(s) can produce the power required to achieve flight manual performance.
This check does not have to be done at maximum engine power. Figure 58-6 is a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the rotorcraft flight manual performance. Some power assurance curves have maximum allowable No limits that must not be exceeded for a given torque value. An in-flight power assurance check may be used in addition to the pretakeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated power values. This occurs when the curve of corrected power and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished.
(6) Deteriorated Engine Power - Turboshaft Engine.
Background.
(A) A specific engine model may have been certificated for operation with power which has "normally" deteriorated below specification. This "normal" deterioration refers to a gradual loss in engine performance, possibly caused by compressor erosion, as opposed to a sudden performance loss which may be due to mechanical damage. The application for deteriorated engine power should not be confused with the installed mechanical engine derating which is frequently used to match transmission and engine power capabilities.
(B) The use of deteriorated power is intended to allow continued operations with an engine which is serviceable and structurally sound, although aircraft performance may be depreciated. The useful life of the engine may, therefore, be extended at a dollar savings to the operator.
(C) Although installed performance is the primary topic in this discussion, considerations must be given to other operational characteristics and systems which may be affected by deteriorated engine power. These include:
(1) Engine characteristics (§ 27.939). Surge margin, engine response, and air-restart capability might be affected and should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism.
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