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. OE 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.
Page B - 15
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 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 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 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.
(1) 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 tuel 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.
Page B - 17
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 OE power ratings. For rotorcraft approved for the use of 30-second/2-minute OEl, 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 OE 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.
Page B - 18
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.
Page B - 19
AC 27-1B, Chg 4 3/31/2014
CERTIFICATION BASIS 14 CFR part 27 CAR 6 CAR 06 6-Amdt.
Ramts 27-Amdt. 21 27-Amdt. 2
1. MGW 1. MGW Sea Sea Level H-V 2. Max. 2. Wt. selected Ref. OGE by applicant at
27.25 wt. Lesser lessor of:
27.1519 27.1587 27.79 Conditions Test of: a. Max. alt. cap. b. 7000' Hd a. Max. alt. cap.
6.116 b. 7000' Hd
6.741
6.743 Hd Hd
3. H-V is 3. H-V is perf. 3. H-V is 3. H-V is 3. H-V is 3. H-V is H-V is perf. Info. info. perf. info. perf. info. perf. info. operating operating
4. Max. alt. 4. If H-V wt is limitation. limitation.
for which less than IGE REM H-V is valid. wt., wt. becomes limitation for
t.o./ldg.
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
CERTIFICATION BASIS 14 CFR part 27 CAR 06 Ramts 27-Amdt. 21 Original Original Original
1. MGW Sea 1. MGW Sea 1. MGW Sea 1. No specific Level Level Level requirement.
IGE 2. Max. IGE 2. Wt. 2. Max.
CONTROL wt. Lesser of: selected Dy approved wt.
Ref. a. Max. alt. applicant to for t.o./ldg at
27.25 cap. max. t.o. and alt above sea
27.1587 27.143 Conditions Test 3. Critical CG 4. Critical b. 7000' Hd Idg. alt. 3. Critical CG level. 3. Critical CG AC 27-1B, Rotor RPM 4. Critical 4. Critical Sections 5. Wind of Rotor RPM Rotor RPM
27.45 & not less than 5. Critical wt. 5. Wind not
27.143 17 kts. 6. Wind of less than 20 not less than mph.
20 mph
6. Max. safe 7. Max. safe 6. Max. sate REM wind is perf. wind is perf. wind is perf.
info. info. info.
7. If 17 kts.
wind demo'd to alt. less than 7000', a Remarks corresponding WAT limit must be established.
FIGURE AC 27.45-2 IGE CONTROLLABILITY REQUIREMENTS Page B - 21
AC 27-1B, Chg 4 3/31/2014 500 500 CORRECTED SHAFT HORSEPOWER ENG SIN 12345 400 300 MIN CERTIFIED
SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500
CORRECTED TOT F FIGURE AC 27.45-3 SHAFT HORSEPOWER VS TURBINE OUTLET
TEMPERATURE - SEA LEVEL STANDARD DAY Page B - 22
16 PRESSURE ALTITUDE - FEET (X 1000) 300 400 500
SHAFT HORSEPOWER AVAILABLE FIGURE AC 27.45-4 UNINSTALLED TAKEOFF POWER AVAILABLE Page B - 23
AC 27-1B, Chg 4 3/31/2014 jo 10 = 0 10°0 L -20 ° C ·0 = 10=501C
TRANSMISSION LIMIT PRESSURE ALTITUDE - FEET (X 1000)
• 200 300 400
SHAFT HORSEPOWER AVAILABLE FIGURE AC 27.45-5 INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Page B - 24
CONDITIONS: EXAMPLE. 3/31/2014 ZERO AIRSPEED OAT = 17.5°0 RPM 100% TOT = 760 ° 0 FIGURE AC 27 45-6 POWER ASSURANCE CHECK CHART ANTH-ICE OFF GENERATOR LOAD 100 AMPS Hp = 4000 FT REQUIRED TORQUE = 70% 1000 FT 800°C B000 FT 780°C 6000 FT 760°C 4000 FT 740°C 2000 FT 720°C SEA LEVEL 700°C 680°C
• 660°C 20 30 40 60 60 80 100 120 OUTSIDE AIR TEMPERATURE - °C PERCENT TORQUE Page B - 24.1 AC 27-1B, Chg 4