SUBPART B - FLIGHT
PERFORMANCE AC 27.45. § 27.45 (Amendment 27-14) PERFORMANCE - GENERAL.
a. Explanation.
(1) Section 27.45 of 14 CFR regulations 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 acceptable test conditions. The test results should be analyzed and expanded by an approved methodology. The guidance within this section is considered an approved methodology.
(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 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.
(4) The 30-second and 2-minute OEI power ratings, defined in 14 CFR 1.1, are based on up to three periods of use during a single flight. The purpose of the three periods 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 the use of these time-limited power ratings is only permitted once in each of the above three periods (i.e., 30-second power must not be used more than once during the initial recovery from an engine failure).
(5) All engines operating (AEO) performance must be based upon approved AEO power ratings. OEI power ratings cannot be applied to an AEO condition.
b. Procedures.
(1) Winds for Testing.
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AC 27-1B, Chg 4 3/31/2014
(i) 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 accounted for, hover performance testing should be conducted in winds of 3 knots or less.
(ii) IGE 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.
(iii) 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) HV - 0 to 5 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: Extrapolation and Interpolation.
(i) Using FAA//AUTHORITY approved methodology:
(A) Hover performance may be extrapolated from test data up to a maximum of ±4,000 feet density altitude from the test altitude.
(B) IGE handling qualities, height-velocity, and engine operating characteristics may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude.
(C) Cruise stability and controllability tests should be evaluated at a minimum 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 density altitude.
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(ii) As in all testing, extrapolation 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. 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.
(A) 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 enroute altitude.
(B) Maximum takeoff and landing altitude may be limited by the IGE hover ceiling for a rotorcraft as described in § 27.73 (through Amendment 27-14) or by low speed controllability. The hover ceiling and any information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certificated to CAR 6, Amendment 6-7 or any amendment of part 27, a hover ceiling may not be presented above the altitude at which takeoff, HV, and IGE controllability tests were conducted plus allowable extrapolation, unless that extrapolated altitude is to at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, including allowable extrapolation, 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.
(ii) 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).
(5) HV envelope (§ 27.79).
(6) IGE controllability (§ 27.143(c)).
(7) Cooling (§§ 27.1041, 27.1043, and 27.1045).
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AC 27-1B, Chg 4 3/31/2014
(8) Engine operating characteristics (§ 27.939).
Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this guidance section.
(B) As detailed in paragraph b.(2) above, the maximum allowable extrapolation of HV, 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.
(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 AC 27.45-1 and AC 27.45-2.
(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 paragraph b.(3)(ii)(A) above.
(4) Temperature Effects.
(i) 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 materials. Therefore, to satisfy § 27.1309(a), the Page B - 12
applicant needs to substantiate the total rotorcraft throughout the foreseeable range of operating temperatures.
(ii) 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/VD evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.
(B) Extrapolation or interpolation of test data should only be allowed if the applicant’s predicted or calculated data is verified by actual test. Extrapolations or interpolations should not exceed 10°C below or 20°C above those values tested.
(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 speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specified 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 AC 27.45-3 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 power Page B - 13
AC 27-1B, Chg 4 3/31/2014 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 AC 27.45-4.
(E) Several installation losses (i.e., power decrements) 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 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 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 AC 27.45-5 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.
(ii) 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 Page B - 14
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 or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM 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 demand to takeoff 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 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. OEI 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.
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AC 27-1B, Chg 4 3/31/2014 (E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical NG 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 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 AC 27.45-6 is a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the RFM performance. Some power assurance curves have maximum allowable NG 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.
(i) 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 Page B - 16
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 or aircraft installation and fuel scheduling mechanism.
(2) Performance of customer bleed air systems may be degraded slightly.
No problem would be anticipated unless certain items within the system depend on a critical engine bleed air pressure for their function.
(3) The maximum attainable gas producer speed, and thus power available under certain ambients, may be affected if engine bleed air pressure is an input to the fuel scheduling mechanism.
(4) Systems for surge protection which schedule on engine bleed air pressure such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The effect would normally be negligible unless when installed, the installation losses, combined with reduced engine bleed air pressure because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.
(ii) Procedures.
(A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in paragraph b.(5) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines is greatly diminished and perhaps eliminated.
(B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in paragraph b.(5) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur with cold temperatures and high altitudes; bleed valves and other devices which schedule on gas producer discharge pressure are most likely to function and reduce power on a hot day at low altitude.
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AC 27-1B, Chg 4 3/31/2014 (C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated engine monitoring procedures will be introduced when allowing the use of deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in- flight procedure must be evaluated by flight tests as described in paragraph b.(5) above. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure already in place for the specification engine, which could eliminate the need for flight test evaluation.
AC 27.45A. § 27.45 (Amendment 27-21) GENERAL.
a. Explanation. Amendment 27-21 adds § 27.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered rotorcraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the RFM.
b. Procedures. All of the guidance material pertaining to AC section 27.45 remains in effect. In addition, the power assurance information included in the RFM should be verified. Although this requirement is normally met with a power assurance curve, other methods of compliance may be proposed.
AC 27.45B. § 27.45 (Amendment 27-21) GENERAL.
a. Explanation. Although § 27.45 was not changed by Amendment 27-29, that amendment added requirements for certification of 30-second/2-minute OEI power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level.
b. Procedures. MG 9 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved. All of the guidance material pertaining to sections 27.45 and 27.45A of this AC remain in effect.
AC 27.45C. § 27.45 (Amendment 27-21) PERFORMANCE - GENERAL.
a. Explanation. Although § 27.45 was not changed by Amendment 27-44, that amendment added new performance and handling qualities requirements for normal category rotorcraft. Included within these regulatory changes is OGE handling qualities.
Additionally, hover performance requirements were re-identified from § 27.73 to § 27.49; HV envelope requirements were re-identified from § 27.79 to § 27.87.
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b. Procedures. All of the guidance material pertaining to AC sections 27.45, 27.45A, and 27.45B remain in effect. In addition, the following apply:
(1) OGE handling qualities may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude.
(2) Hover performance guidance that applied to § 27.73 is applicable to § 27.49.
(3) HV envelope guidance that applied to § 27.79 is applicable to § 27.87.
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AC 27-1B, Chg 4 3/31/2014
CERTIFICATION BASIS
14 CFR part 27 CAR 6 CAR 06 Rqmts 27-Amdt. 21 27-Amdt. 2 Original 6-Amdt. 7 6-Amdt.
4 Original Original H-V Ref.
27.25
27.79
27.1519
27.1587 Test Conditions
1. MGW Sea Level
2. Max.
OGE wt. Lesser of:
a. Max. alt.
cap.
1. MGW Sea Level
2. Wt. selected by applicant at lessor of:
a. Max. alt. cap.
b. 7000' Hd
1. MGW Sea Level
2. Wt.
selected by applicant at lessor of:
1. MGW Sea Level
2. Wt.
selected by applicant at lessor of:
1. MGW Sea Level
2. Wt.
selected by applicant max. t.o.
1. MGW Sea Level
2. Wt.
selected by applicant at
1. No specific wt. and alt.
requirements.
6.116
6.741
6.743
b. 7000' Hd a. Max.
alt. cap.
b. 7000' Hd
a. Max.
alt. cap.
b. 7000' Hd and ldg. alt.
max. t.o.
and ldg. alt.
RFM
3. H-V is perf. Info.
4. Max. alt.
for which H-V is valid.
3. H-V is perf.
info.
4. If H-V wt is less than IGE wt., wt. becomes limitation for
t.o./ldg.
3. H-V is perf. info.
3. H-V is perf. info.
3. H-V is perf. info.
3. H-V is operating limitation.
H-V is operating limitation.
AC 27-1B 5. If H-V is Section less than
27.45 & OGE wt., H-
27.51 V wt.
becomes limit.
6. Applicant is encouraged to demo H-V Remarks to WAT limits.
7. Hover data may be shown above 7000' if H-V & IGE are demo'd to 7000'.
FIGURE AC 27.45-1 HV Requirements Page B - 20