Page B - 17 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. An engine may be certified to produce a specification horsepower at a particular temperature or engine speed rating with higher maximum limit value approved. Only the rating values should be used to determine the installation losses.
The limit values of engine temperature and/or speed are established and certified to allow specification powers to continue to be developed as the engine deteriorates in service.
(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 installed specification engine installed on a production rotorcraft is capable of producing the horsepower 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 a total of the different minimum installed specification horsepower values. 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, 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 NG speed to a constant corrected speed at cold temperatures. Others may limit power to a minimum 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, and 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 29.45-6 is a typical power assurance curve for an installed engine showing minimum acceptable
Page B - 18 torque which assures that power is available to meet the rotorcraft flight manual 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 pre-takeoff 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 horsepower values. This occurs when the curve of measured corrected horsepower 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.
(7) Deteriorated Engine Power - Turboshaft Engine.
(i) Background.
(A) A specific engine model ma y 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 depreciated engine power. These include:
( 1 ) Engine characteristics (§ 29.939). The reduced compressor discharge pressure, PC, would reduce engine surge margin and possibly affect engine response and engine air-restart capability. These items should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism.
( 2 ) Performance of customer bleed ai r systems may be degraded slightly.
No problem would be anticipated unless certain items within the system depend on a critical PC for their function.
Page B - 19 ( 3 ) The maximum attainable gas producers speed, and thus power available under certain ambients, may be affected if PC pressure is an input to the fuel scheduling mechanism.
( 4 ) Systems for surge protection which schedule on P C pressure such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The affect would normally be negligible unless when installed, the installation losses combined with reduced PC 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 (6) (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 testi ng to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in paragraph (6) (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 cold and high; 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.
(C) The adjustments to the norma l 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 region established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated procedures will be introduced for 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 (6) (engine power-turboshaft engines) herein. 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.
(D) If a complex power assuranc e procedure is presented with involved data reduction and trending requirements, consideration should be given to restricting the use of deteriorated power to operators where close control over operations is
Page B - 20 exercised and/or the operator has demonstrated his ability to operate safely with deteriorated engines.
AC 29.45A. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation . Amendment 29-24 adds § 29.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered aircraft. 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 rotorcraft flight manual (RFM).
b. Procedures . All of the policy material pertaining to this section 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 29.45B. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation . Amendment 29-34 added the requirements for certification of 30-second/2-minute One Engine Inoperative (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 . Information provided in paragraph AC 29 MG 9 includes guidance material on power assurance procedures to ensure that the OEI power level can be achieved. The guidance material presented in paragraphs AC 29.45 and AC 29.45A continue to apply.
Page B - 21
CERTIFICATION BASIS
FAR 29 CAR 7 Requirement s 29-Amdt. 21 29-Amdt. 1 Original Original H-V Ref.
29.25
29.87
29.1517
29.1581
29.1583
7.11
7.715
7.741 CAT A TEST
CONDITIONS Cat A & B (>9 pax seats):
1. W.A.T. for which
t.o. and ldg. are approved.
2. Failure of critical engine.
Cat A:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. capability.
3. Failure of critical engine.
Cat A:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. capability.
3. Failure of critical engine.
Cat. A:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. Capability.
3. Failure of critical engine.
AC 29-2C paras AC
29.45 & AC
29.79 CAT A RFM
3. H-V is limitation.
4. Type of ldg.
surface.
4. H-V is limitation.
5. Type of ldg.
Surface.
4. H-V is limitation.
5. Type of ldg.
surface.
4. H-V is limitation.
5. Type of ldg.
surface.
CAT B TEST
CONDITIONS Cat B (</=9 pax seats):
1. MGW Sea Level
2. Max. OGE wt.
Max.
alt. Capability
3. Complete power failure, or failure of critical engine (w/eng isolation).
Cat B:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. Capability
3. Complete power failure, or failure of critical engine (w/eng isolation).
Cat B:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. Capability
3. Complete power failure, or failure of critical engine (w/eng isolation.)
Cat B:
1. MGW Sea Level
2. Max. IGE wt.
Max.
alt. Capability
3. Complete power failure, or failure of critical engine (w/eng isolation.)
CAT B RFM
4. H-V is perf. Info.
5. Type of ldg.
surface.
4. H-V is limitation info.
5. Type of ldg.
Surface.
4. H-V is limitation info.
5. Type of ldg.
surface.
4. H-V is limitation info.
5. Type of ldg.
surface.
FIGURE AC 29.45-1. H-V REQUIREMENTS
Page B - 22
CERTIFICATION BASIS
FAR 29 CAR 7 Requirement s 29-Amdt. 24 29-Amdt. 3 Original Original IGE
CONTROL Ref. 29.25
29.1583
7.121
7.743
CAT A TEST
CONDITIONS Cat A
1. W.A.T. for which
t.o. and ldg. are approved.
2. Critical wt.
Critical CG Critical Nr
3. Wind not less than 17 kts.
1. Conditions selected by the applicant.
2. Critical CG Critical Nr
3. Wind not less than 17 kts.
1. Conditions selected by the applicant.
2. Critical CG Critical Nr
3. Wind not less than 20 mph.
1. Conditions selected by the applicant.
2. Critical CG Critical NR
3. Wind not less than 20 mph.
AC 29-2C paras AC
29.45 and AC 29.143 CAT A RFM
4. Max.
allowable wind is limitation.
4. Max safe wind above max.
alt. For which 17 kt. Wind envelope is established is perf. info.
4. Max safe wind above alt. for which 17 kt. wind envelope is established is perf. info.
4. Max.
allowable wind above the altitude for which 20 mph wind envelope is est. is perf.
info.
CAT B TEST
CONDITIONS Cat B:
1. W.A.T. for which
t.o. and ldg. are approved.
2. Critical wt.
Critical CG Critical Nr
3. Wind speed & quad selected by the applicant.
CAT B RFM
4. Max. safe wind is perf.
info.
FIGURE AC 29.45-2 IGE CONTROLLABILITY REQUIREMENTS
600 500 CORRECTED SHAFT HORSEPOWER ENG S/N 12345
MIN CERTIFIED
SPECIFICATION ENGINE 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE AC 29.45-3 SHAFT HORSEPOWER VS TURBINE OUTLET
TEMPERATURE - SEA LEVEL STANDARD DAY Page B - 23
20 16 PRESSURE ALTITUDE - FEET (X 1000) 200 ЗАМІ 500
SHAFT HORSEPOWER AVAILABLE FIGURE AC 29.45-4 UNINSTALLED TAKEOFF POWER AVAILABLE Page B - 24
20 18 70= 10-501 C 14 TRANSMISSION LIMIT PRESSURE ALTITUDE - FEET (X 1000) 200 300 400
SHAFT HORSEPOWER AVAILABLE FIGURE AC 29.45-5 INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Page B - 25
120 EXAMPLE: OAT = 17.5°C TOT = 760° C Hp = 4000 FT REQUIRED TORQUE = 70% 80 100 PERCENT TORQUE FIGURE 29.45-6 POWER ASSURANCE CHECK CHART GENERATOR LOAD 100 AMPS 60 OUTSIDE AIR TEMPERATURE - °C CONDITIONS: ZERO AIRSPEED RPM 100% ANTI-ICE OFF 20 30 0 10 800°C 780°C 760°C 740°C 720°C 700°C 680°C 660°C Page B - 26
Page B - 27 AC 29.49. § 29.49 (Amendm ent 29-39) PERFORMANCE AT MINIMUM OPERATING SPEED. HOVER PERFORMANCE FOR ROTORCRAFT.
(For performance at minimum operating speed and for hover performance prior to Amendment 39, see § 29.73 and paragraph AC 29.73.)
a. Explanation .
(1) Amendment 29-39 redesignated § 29.73 as § 29.49 to relocate the requirements for rotorcraft hover performance. For the purpose of this manual, the word “hover” applies to a rotorcraft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air.
(2) Under § 29.49, hover perform ance should be determined at a height consistent with the takeoff procedure for Category A rotorcraft and IGE for Category B rotorcraft. Additionally, OGE hover performance should be determined for both Category A and B rotorcraft. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ rotor diameters above the surface.
(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 nondimensional power coefficients (Cp) and thrust coefficients (Ct) for normalizing and presenting test results, a minimum amount of data are required to cover the rotorcraft’s performance 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 landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.
(5) The minimum hover height 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 paragraph AC 29.51 for the procedure to determine the minimum allowable hover height.
b. Procedures .
(1) Two methods of acquiring hover per formance data are the tethered and 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
Page B - 28 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 Ct/Cp spread. The load cell reading is recorded for each stabilized point. The total thrust the rotor produces is the rotorcraft’s gross weight, 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.
(2) The pilot maintains the rotorcraft in position so that the cable and load cell are perpendicular to the ground. To insure 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. This can also be accomplished by attaching two accelerometers to the load cell which sense 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. Accurate load cell values may also be obtained by measuring cable angles and, through geometry, determining a corrected load cell value. 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 wi nd velocity should be 3 knots or less.
Large rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are:
(i) Engine torques.
(ii) Rotor speed.
(iii) Ambient temperatures.
(iv) Pressure altitude.
(v) Fuel used (or remaining).
(vi) Load cell reading.
(vii) Generator(s) load.