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
AC 27-1B, Chg 4 9/17/2009 AC 27.49. § 27.49 (Former § 27.73) (Amendment 27-44) Performance at Minimum Operating Speed.
(For § 27.73 prior to Amendment 27-44, see AC 27.73)
a. Explanation. Amendment 27-44 adds a requirement to determine out-of-groundeffect (OGE) hover performance. Once reserved for special missions, OGE operations are now a common practice.
(1) The word "hover" applies to a helicopter that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air. For the purpose of this manual, the word "hover" will mean pure hover.
(2) The regulatory requirement for hover performance, § 27.49, refers to hover in-ground-effect (IGE) and OGE. Hover OGE is the absence of measurable ground effect. Hover OGE is established when the power required to hover is the same at different heights above the ground.
(3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes.
Using non-dimensional power coefficients (Cp) and thrust coefficients (Ct) for normalizing and presenting test results minimizes the amount of data required to cover the helicopter's operating envelope.
(4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landing. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.
(5) The IGE hover ceiling for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to AC 27.51 for the procedure to determine this hover height.
b. Procedures.
(1) Two methods of acquiring hover performance data are the tethered and the free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft's pull on the cable. Hover heights are based on skid or wheel height above the ground. During tethered hover tests, the rotorcraft should be at light gross weight. The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This technique will produce a large COp spread. The load cell reading is recorded for each Page B - 24 2
stabilized point. The total thrust the rotor produces is equal to the rotorcraft's gross weight plus the weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For some rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data. IGE hover performance may be affected by the composition and slope of the surface beneath the rotorcraft. Therefore, the tests should be conducted over a smooth, level, hard surface.
(2) The pilot maintains the rotorcraft in position so that the cables and load cell are perpendicular to the ground. To ensure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. Positioning can also be accomplished by attaching two accelerometers to the load cell, which sense angle or movement along the longitudinal and lateral axes.
Any displacement of the load cell will be reflected on instrumentation in the cockpit, and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft's internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency.
(3) To obtain consistent data, the wind velocity should be less than 3 knots as there are no accurate methods of correcting hover data for wind effects. To minimize inaccuracies due to hysteresis, collective movement should be made in only one direction. Rotorcraft with high downwash velocities may tolerate higher wind velocities.
The parameters usually recorded at each stabilized condition are:
(i) Engine and transmission torque.
(il) Rotor speed.
(ili) Ambient and engine temperatures, such as Measured Gas Temperature (MGT).
(iv) Pressure altitude.
(v) Fuel used (or remaining).
(vi) Load cell reading.
(vii) Generator(s) load.
(vill) Wind speed and direction.
(ix) Hover height.
Page B - 24.3