the use of these time-limited power ratings is only permitted once in each of the above three uses (i.e., 30-second power must not be used more than once during the initial recovery from an engine failure).
(6) 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.
(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 may be tolerated for takeoff and landing distance tests but not for hover performance.
Likewise higher winds may be tolerated during hover performance testing on large, heavy rotorcraft with high rotor downwash velocities than for smaller rotorcraft with lower 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) Past experience has shown that a steady wind of 0 to 10 knots will result in acceptable takeoff and landing performance if distances are corrected for the winds measured during these tests. This is not the case for vertical takeoffs and landings. To obtain consistent and repeatable vertical performance data, using wind speeds up to 5 knots is acceptable. In actuality, a rotorcraft may exhibit reduced IGE hover performance in wind speeds from 3 to 15 knots due to partial immersion of the main rotor in its own vortex. Since the height -speed envelope determination is affected by wind just as vertical takeoff and landing performance are, the same allowable winds for testing should be adhered to for HV testing; i.e., 0 to 5 knots. For category A testing, the effects of crosswind and tailwind should also be considered up to the maximum for which category A certification is requested.
(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) Conventional takeoff and landing - 0 to 10 (data to be corrected) (C) Vertical takeoff and landing - 0 to 5 knots (D) HV - 0 to 5 knots Page B – 11
AC 29-2C, Chg 4 5/1/2014
(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests. The wind effects on required runway length for takeoff and landing distances may be shown in the flight manual.
(v) Full wind credit may be given for conventional takeoff and landing field lengths. This credit should not be more than the nominal wind component along the takeoff or landing path opposite to the direction of flight.
(2) Altitude Effects: Extrapolation and Interpolation.
(i) Using FAA/AUTHORITY approved methodology:
(A) Hover, takeoff, and landing performance may be extrapolated from test data up to a maximum of ±4,000 feet density altitude from the test altitude.
(B) Experience has shown that 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 of 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.
(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.
(i) 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 § 29.1527 and delineates the maximum altitude up to which operation is allowed. This altitude normally constitutes the maximum cruise or enroute altitude.
(B) Maximum weight, altitude, and temperature for takeoff and landing constitute a limitation. The maximum takeoff and landing altitude may be coincident with but never above the maximum operating altitude limitation. Takeoff and landing, hover ceiling data, and presentation requirements are presented in §§ 29.51, 29.53,
29.59, 29.63, 29.73, 29.1583, and 29.1587.
Page B – 12
(ii) Procedures.
(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required:
(1) Takeoff (§§ 29.51-29.63)
(2) Climb (§§ 29.64-29.67)
(3) Performance at minimum operating speed (§ 29.49)
(4) Landing (§ 29.75)
(5) HV envelope (§ 29.87)
(6) IGE controllability (§ 29.143(c))
(7) Cooling (§§ 29.1041-29.1045)
(8) Engine operating characteristics (§ 29.939) 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.
(C) Prior to Amendment 29-21, HV information was an operating limitation.
With the adoption of Amendment 29-21, the HV curve is performance information for category B rotorcraft with nine or less passenger seats but remains a limitation for category A rotorcraft and category B rotorcraft with 10 or more passenger seats.
(D) Prior to Amendment 29-24, IGE controllability was required in 17 knots of wind to the maximum takeoff and landing conditions. With the adoption of Amendment 29-24, if IGE or OGE hover performance is presented for a category B rotorcraft to an altitude in excess of that for which IGE controllability at 17 knots is presented, the maximum safe wind demonstrated for hover operations must be presented in the RFM. The amendment did not change the requirement for category A rotorcraft.
(E) 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 29.45-1 and AC 29.45-2.
Page B – 13
AC 29-2C, Chg 4 5/1/2014 (F) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in paragraph b.(2) above and not above the lowest limiting altitude resulting from the requirements listed in paragraph b.(3)(ii)(A) above.
(4) Temperature Effects.
(i) Background.
(A) The regulations prohibit any unsafe design feature throughout the range of environmental conditions for which certification is requested. The regulations also require that the performance and handling qualities be determined over the approved range of atmospheric variables selected by the applicant.
(B) Substantiation of temperature effects on performance and handling characteristics is required throughout the approved temperature range. 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 blade 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.
(C) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected, sometimes significantly, in both the performance and flying qualities areas. Hot temperature conditions although not shown to be as critical should be given consideration.
(D) 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 § 29.1309(a), the applicant needs to substantiate the total rotorcraft at the extreme temperatures for which certification is requested.
(ii) Procedures.
(A) The FAA/AUTHORITY is responsible for verifying the applicant’s predictions of performance and handling characteristics at the temperature extremes for which certification is requested. A limited flight verification, if necessary, could include spot checks of hover and climb performance, IGE controllability, roughness determination, simulated power failure, static stability, height-velocity, VNE/VD evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.
Page B – 14
(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) Weight Effects. Test weights should be maintained within +3 percent and -1 percent of the target weight for each data point. Weight may be extrapolated or interpolated only along an established W/σ line within the allowable altitude extrapolated range.
(6) 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 horsepower used to complete the flight manual performance must be based on horsepower 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 horsepower values at specification temperatures and speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specified horsepower 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 horsepower. 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 29.45-3 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 horsepower values 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 Page B – 15
AC 29-2C, Chg 4 5/1/2014 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 29.45-4.
(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 horsepower resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet 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 AC 29.45-5 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.
(ii) 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 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 is held constant. This is a characteristic Page B – 16
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, or 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 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 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 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. 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, etc., must be accounted for accordingly.
Page B – 17
AC 29-2C, Chg 4 5/1/2014 (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 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 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 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 Page B – 18
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 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 producers 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 affect 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.(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 b.(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 Page B – 19
AC 29-2C, Chg 4 5/1/2014 discharge pressure are most likely to function 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. 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 b.(6) 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.
(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 or the operator has demonstrated the ability to operate safely with deteriorated engines.
(8) Engine Failure Testing Considerations
(i) For all tests to examine behavior following an engine failure, usually the failure of the engine is simulated in some way. For engines with a hydro-mechanical governing system, it is common practice to close the throttle quickly to idle. For rotorcraft equipped with a FADEC, and particularly those with a 2 minute/30 second OEI rating structure, it is common practice to simulate an OEI condition by using reduced power on all engines by means of a flight test tool.
(ii) In every case, it must be demonstrated that all aspects of rotorcraft and powerplant behavior are identical to those that would occur in the event of an actual engine failure with the remaining engine developing minimum-specification power. Of particular concern are “dead engine” power decay characteristics, “live engine” acceleration characteristics, and rotor RPM control.
(iii) To this end, it is expected that a number of actual engine shut down tests will be conducted to generate sufficient data to validate the fidelity of the flight test tool and methodology, which will then allow its use in developing regulatory performance data. In general, it is best to conduct the tests in a low hover with the rotorcraft stabilized below the HV low point. An engine is then shut down and, following the appropriate pilot intervention time, the collective is raised to cushion the landing.
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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 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 29.45B. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation. Although § 29.45 was not changed by Amendment 29-34, 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 material on power assurance procedures to ensure that the OEI power level can be achieved. All of the guidance material pertaining to AC sections 29.45 and 29.45A remain in effect.
AC 29.45C. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation. Although § 29.45 was not changed by Amendment 29-51, that amendment added new performance and handling qualities requirements for transport category rotorcraft. Included within these regulatory changes is OGE handling qualities.
Additionally, hover performance requirements were re-identified from § 29.73 to § 29.49.
b. Procedures. All of the guidance material pertaining to AC sections 29.45,
29.45A, and 29.45B 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 § 29.73 is applicable to § 29.49.
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AC 29-2C, Chg 4 5/1/2014
CERTIFICATION BASIS 14 CFR pa rt 29 CAR 7 Rqms 29-Amdt. 21 29-Amdt. 1 Original Original HV CAT A Cat A & B (>9 Cat A: Cat A: Cat. A:
Ref. TEST pax seats): 1. MGW Sea 1. MGW Sea 1. MGW Sea
29.25 CONDITIONS 1. W.A.T. for Level Level Level
29.87 which t.o. and 2. Max. IGE 2. Max. IGE 2. Max. IGE wt.
29.1517 ldg. are wt. Max. wt. Max. Max.
29.1581 approved. alt. capability. alt. capability. alt. Capability.
29.1583 2. Failure of 3. Failure of 3. Failure of 3. Failure of
7.11 critical engine. critical engine. critical critical engine.
7.715 engine.
7.741 AC 29- CAT A 3. HV is 4. HV is 4. HV is 4. HV is 2C, RFM limitation. limitation. limitation. limitation.
Sections 4. Type of ldg. 5. Type of ldg. 5. Type of ldg. 5. Type of ldg.
29.45 & surface. Surface. surface. surface.
29.79 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 4. HV is perf. 4. HV is 4. HV is 4. HV is RFM Info.
5. Type of ldg.
surface.
limitation info.
5. Type of ldg.
Surface.
limitation info.
5. Type of ldg.
surface.
limitation info.
5. Type of ldg.
surface.
FIGURE AC 29.45-1. HV REQUIREMENTS Page B – 22