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 horsepower 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 usually conducts test to confirm the installed specification. Methods used vary widely between manufacturers, but usually include some combination of ground and flight tests.
Figure 55-8 is a typical example of an installed power available chart for one set of conditions.
(F) This predicted installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion decreases the available horsepower output and 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 horsepower available from a minimum installed specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual horsepower available from a minimum installed specification engine at the published climb airspeed.
This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.
Procedures.
(A) To this point the minimum installed specification engine horsepower 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 an initial application of power, 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 is 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 applicable value (takeoff, 30-second, and 2 1/2-minute power), the engine is first stabilized at a low power setting. After stabilization, rapidly increase the power demand to takeoff, 30-second and 2 1/2-minute power levels as necessary. Record the engine parameters as soon as the specification 70 Par 55
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 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 1/2-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. 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 No 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.
Par 55 71
(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 55-9 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 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 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 depreciated engine power. These include:
72 Par 55
(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 air systems may be degraded slightly. No problem would be anticipated unless certain items within the system depend on a critical Pc for their function.
(3) The maximum attainable gas producers speed, and thus power available under certain ambients, may be affected if Po pressure is an input to the fuel scheduling mechanism.
(4) Systems for surge protection which schedule on Pc 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.
(її) 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 testing 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 tunction and reduce power on a hot day at low altitude.
(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 region established as a condition for the engine to remain in service when below specification.
Par 55 73
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 assurance 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 exercised and/or the operator has demonstrated his ability to operate safely with deteriorated engines.
55A. § 29.45 (Amendment 29-24) 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.
55B. § 29.45 (Amendment 29-24) 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 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. Information provided in Appendix 2 of this AC includes guidance material on power assurance procedures to ensure that the OEI power level can be achieved. The guidance material presented in Paragraphs 55 and 55A continue to apply.
74 Par 55
CERTIFICATION BASIS FAR 29 CAR 7 Requirements 29-Amdt. 21 29-Amdt. 1 Original Original 7/30/97 Cat A & B (>9 pax Cat A: Cat A: Cat. A:
H-V Ref. seats): 1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level
29.25 1. W.A.T. for which 2. Max. IGE wt. Max. 2. Max. IGE wt. Max. 2. Max. IGE wt. Max.
29.87 t.o. and Idg. are alt. capability. alt. capability. alt. capability.
29.1517 CAT A TEST approved. 2. Failure of critical engine. 3. Failure of critical 3. Failure of critical engine. 3. Failure of critical engine.
29.1581 CONDITIONS engine.
29.1583
7.11
7.715
7.741
3. H-V is limitation. 4. H-V is limitation. 4. H-V is limitation. 4. H-V is limitation.
AC 29-2A CAT A 4. Type of Idg. 5. Type of Idg. 5. Type of Idg. surface. 5. Type of Idg.
Paragraphs RFM surface. surface. surface.
55 & 72 Cat B (</=9 pax seats): Cat B: Cat B: Cat B:
1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level
2. Max. OGE wt. Max. 2. Max. IGE wt. Max. 2. Max. IGE wt. Max. 2. Max. IGE wt. Max.
CAT B alt. Capability alt. Capability alt. Capability alt. Capability TEST 3. Complete power 3. Complete power 3. Complete power 3. Complete power CONDITIONS failure, or failure of failure, or failure of failure, or failure of failure, or failure of critical engine (w/eng critical engine (w/eng critical engine (w/eng critical engine (w/eng isolation). isolation). isolation.) isolation.)
4. H-V is perf. Info. 4. H-V is limitation 4. H-V is limitation info. 4. H-V is limitation CAT B 5. Type of Idg. info. 5. Type of Idg. surface. info.
RFM surface. 5. Type of Idg. 5. Type of Idg.
surface. surface.
FIGURE 55-1. H-V REQUIREMENTS
CERTIFICATION BASIS FAR 29 CAR 7 Requirements 29-Amdt. 24 29-Amdt. 3 Original Original Cat A Conditions 1. Conditions 1. Conditions
1. W.A.T. for which selected by the selected by the selected by the IGE t.o. and Idg. are applicant. applicant. applicant.
CONTROL approved. 2. Critical CG 2. Critical CG 2. Critical CG Ref. 29.25 CAT A TEST 2. Critical wt. Critical Nr Critical Nr Critical NR
29.1583 CONDITIONS Critical CG 3. Wind not less 3. Wind not less 3. Wind not less
7.121 Critical Nr than 17 kts. than 20 mph. than 20 mph.
7.743 3. Wind not less than 17 kts.
4. Max. allowable 4. Max safe wind 4. Max safe wind 4. Max.
wind is limitation. above max. alt. for above alt. for which allowable wind which 17 kt. wind 17 kt. wind envelope above the Paragraphs 55 AC 29-2A and 80 CAT A RFM envelope is established is perf. is established is perf. info. altitude for which 20 mph wind info. envelope is est.
is perf. info.
Cat B:
1. W.A.T. for which
t.o. and Idg. are approved.
CAT B 2. Critical wt.
TEST Critical CG CONDITIONS Critical Nr
3. Wind speed & quad selected by the applicant.
CAT B 4. Max. safe wind is 7/30/97 RFM perf. info.
FIGURE 55-2. IGE CONTROLLABILITY REQUIREMENTS
600 500 CORRECTED SHAFT HORSEPOWER ENG S/N 12345 400 300
MIN CERTIFIED
SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE 55-3. SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE -
SEA LEVEL STANDARD DAY 77
20 | 18 -30° to -50° 16 PRESSURE ALTITUDE - FEET (X 1000) 2 200 300 400 500
SHAFT HORSEPOWER AVAILABLE FIGURE 55-4. UNINSTALLED TAKEOFF POWER AVAILABLE 78 Par 55
50 ° C 40 ° C 30 ° C 20 ° C 10 ° C - -10° € -20° C -30 ° 10-50¢C
TRANSMISSION LIMIT PRESSURE ALTITUDE - FEET (X 1000) 200 300 400
SHAFT HORSEPOWER AVAILABLE FIGURE 55-5. INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Par 55 79
CONDITIONS: EXAMPLE: AC 29-2B ZERO AIRSPEED OAT = 17.5°C RPM 100% TOT = 760° C ANTI-ICE OFF Hp = 4000 FT GENERATOR LOAD 100 AMPS REQUIRED TORQUE = 70% 10000 FT 800°C 8000 FT 780°C 6000 FT 760°C 4000 FT 740°C 2000 FT 720°C SEA LEVEL 700°C 680°C 660°C Par 55 OUTSIDE AIR TEMPERATURE - °C 10 20 30 40 60 20 PERCENT TORQUE 40 60 80 100 120 7/30/97 FIGURE 55-6. POWER ASSURANCE CHECK CHART
56. § 29.49 (Amendment 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 69.)
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 performance 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 1½ 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 (Cy) 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 57 for the procedure to determine the minimum allowable hover height.
b. Procedures.
(1) Two methods of acquiring hover performance 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 Par 56 81