flows that would exist at hot day conditions. For example, thermostatically controlled oil cooler flow could be set for hot day conditions.
(3) The component and fluid temperature correction factor to be applied when test ambients do not correspond to the hot day conditions is commonly called the “degree-for-degree correction.” It may be possible to justify, and the regulation allows the application of a more rational, less conservative correction factor. A correction factor other than degree-for-degree should be based on engineering test data.
(4) No corrected temperatures may exceed established limits. In order to maintain temperatures within established limits, the applicant may be willing to accept lesser performance than the full capability of a device. For example, a starter/generator capable of cooling under test cell conditions to 200 amperes continuous load may be limited to a lesser value, perhaps to 150 amperes, when installed in the aircraft due to cooling considerations. This continuous load for cooling must be equal to or greater than the allowable continuous load designated on aircraft instruments.
c. Thermal Limit Correction.
(1) An important correction factor which is not discussed in the regulations, but is frequently necessary to show the cooling adequacy required by § 29.1041, is the thermal limit correction factor. This factor is sometimes required if, at test day conditions, the engine measured gas temperature does not correspond to that which would have occurred on a minimum specification engine at hot day conditions.
(2) The correction factor would not apply to those components not affected by changes in measured gas temperature (MGT) at a constant power. Typical items expected to be affected by changes in the MGT at constant power would be engine oil temperature, thermocouple harnesses, or other fluid, component, or ambient temperatures in the vicinity of the engine hot-section or exhaust gases. Other items remote from the hot-section, perhaps the starter-generator or fuel control, would not be expected to be influenced by MGT variations; however, the items affected and the magnitude of the factor to be applied should be established by testing.
(3) There are several acceptable methods for establishing the appropriate thermal limit correction factor during development testing. The general idea is to establish a stabilized flight condition, typically ground-run or IGE hover, and to vary the measured gas temperature at approximately fixed power and OAT conditions. This may be accomplished by utilizing engine anti-ice bleed air, customer bleed air, or by ingesting warmer than ambient air (either an external source or the engine bleed air) into the engine inlet. Care should be used in ingesting warmer than ambient air to assure that the warm air is diffused in order to avoid possible engine surge.
(i) If it is not possible to attain a suitable variation in MGT by these methods, an acceptable, but more conservative thermal limit correction may be Page E - 141
obtained by allowing both shaft horsepower and MGT to vary at a stabilized flight condition and OAT.
(ii) The component temperature is plotted as a function of MGT, and the thermal limit correction from any test day MGT for any flight condition, to the MGT that would have existed with minimum specification engines on a hot day, is then applied to derive the final measured component temperature.
(4) In certain rare instances, it may not be required that the correction factor be applied to the full thermal limit capability of the engine. Consider the following example for the hot day hover IGE cooling test point at sea level.
Corresponding Power (SHP) MGT (°C) Drive System Limit 900 --- Twin-Engine Hot Day Power Available 1,050 750 Hot Day Power Required at Maximum G.W. 850 650 Engine Maximum Allowable MGT (Instrument Marking) --- 765 Test Day (90° F OAT) Parameters 850 600
(i) Notice that the installed hot day power available MGT from the engine performance program, is 15° C cooler than the limit MGT (750° vs. 765° C), thus the engine has 15° C “field margin” which would allow the engine temperature to gradually increase 15° C to maintain a given power as engine life is utilized. Secondly, the measured gas temperature corresponding to hot day power required at maximum gross weight, is less than that corresponding to either the drive system limit or twin-engine hot day power available. Thus, the thermal limit correction could be applied from the test day MGT, 600° C, to the power required MGT plus the field margin, 650° C plus 15° C, rather than applying the correction factor to the full thermal capability of the engine, 765° C.
(ii) Care should be used in applying this relieving method, because as the hover altitude changes, the maximum gross weight and power required (and the associated MGT) will vary. The data must be corrected to at least the maximum MGT for a minimum specification engine that can occur in service at the flight condition under investigation.
AC 29.1043A. § 29.1043 (Amendment 29-26) COOLING TESTS.
a. Explanation Procedures . Amendment 29- 26 adds a new paragraph to § 29.1043(a)(5), to define “stabilization” as it pertains to powerplant systems cooling tests.
b. . All of the policy material pertaining to this section remains in effect with additional information that “stabilized temperatures” are achieved when the rate of change is less than 2° F per minute.
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AC 29.1045. § 29.1045 CLIMB COOLING TEST PROCEDURES.
a. Objective. The objective of the regulation is to verify, for Category A and for Category B rotorcraft described, that cooling provisions are adequate for a one-engine-inoperative (OEI) climb or descent initiated from a multiengine cruise at the critical altitude with stabilized component temperatures. The specific flight conditions and powers are described in the regulation.
b. Explanation.
(1) This regulation specifies climb or descent cooling with OEI for Category A rotorcraft and for Category B rotorcraft with Category A powerplant isolation and fireproof or isolated structure, controls, etc., which are essential for controlled flight and landing. For the Category B machine described, the testing should be accomplished at the steady rate of climb or descent established under § 29.67(b), i.e., at the best OEI rate of climb (or descent) and the remaining engine at maximum continuous power or 30-minute power, whichever is applicable.
(2) The engine whose shutdown has the most adverse effect on the cooling conditions for the remaining engine(s) and powerplant components should be inoperative.
(3) The regulation provides that the climb cooling test may be conducted in conjunction with the takeoff cooling test of § 29.1047. This possible combining of tests applies only to § 29.1047(a), since § 29.1047(b) is a multiengine climb and not related to the OEI climb procedures of § 29.1045.
c. Procedures.
(1) The OEI climb cooling test point begins from a multiengine cruise, with stabilized fluid and component temperatures, 1,000 feet below either the all-engine-critical altitude or the maximum altitude at which the rate of climb is 150 FPM, whichever is the lowest altitude. If the minimum altitude derived is less than sea level, the climb should begin from a twin engine cruise with stabilized fluid and component temperatures at the minimum practical altitude.
(i) The all-engine-critical altitude is the maximum altitude at which, for the ambient conditions prescribed, it is possible to maintain the multiengine specified power. For example, if for multiengine operations, the transmission maximum continuous torque can be maintained on the hot day to a maximum altitude of 10,000 feet above which power would have to be reduced because of gas temperature or other limitations, then 10,000 feet is the all-engine-critical altitude. Point ”A” in figure AC 29.1045-1 illustrates the all-engine-critical altitude.
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(ii) The 150 FPM climb criteria should be based on multiengine operation at maximum continuous power available at hot day conditions at maximum gross weight.
(iii) Fluid and component temperatures are considered stabilized when the rate of change is less than 2° F per minute.
(2) The OEI climb power to be utilized is 30-minute OEI hot day power available (if approval of 30-minute power on the aircraft is requested), followed by maximum continuous hot day power available. If 30- minute OEI power approval is not requested, the power to be utilized would be maximum continuous hot day power available.
(i) Rotorcraft for which approval of a continuous OEI power rating is requested would use the power available on a hot day at the maximum continuous OEI rating following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).
(ii) If the maximum continuous OEI approval is not requested, then the highest hot day power available approved for continuous usage from the remaining engine(s) under OEI conditions would be used following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).
(3) In order to achieve representative test results, the rotorcraft climb rate and airspeed should approximate those which would occur on a hot day. This is accomplished by adjusting rotorcraft gross weight as required to produce the desired climb rate based on published or predicted climb performance data. The possible adverse effects of climb fuselage attitude on cooling air duct entrances should be considered in the selection of center-of-gravity of the test aircraft.
(4) The OEI climb should be continued for at least 5 minutes after the occurrence of the highest temperature recorded or until the maximum certification altitude is reached. Generally, temperatures would be expected to peak a short time after the climb begins since component and fluid temperatures are stabilized prior to entry to the climb phase.
(5) For Category B rotorcraft, defined in § 29.1045(a)(2) without a positive OEI rate of climb, the descent should begin from a hot day maximum continuous power multiengine cruise, with stabilized fluid and component temperatures, at the all-engine-critical altitude.
(6) The descent should conclude at either the maximum altitude at which level flight can be maintained with one engine inoperative or at the minimum practical altitude, whichever is higher.
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(7) The OEI powers available to be utilized during the descent would be the same as those prescribed previously for OEI climb cooling. OEI operation should continue until component and fluid temperatures stabilize.
(8) The airspeeds utilized in the climb and descents should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeeds most critical for cooling should be used. In no case, however, should it be required that the selected airspeeds exceed the speeds established under §§ 29.67(a)(2) and 29.67(b).
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MULTIENGINE POWER AVAILABLE 20K HOT DAY — 15K — PRESSURE ALTITUDE OAT 10° C 10K - - CRITICAL ALTITUDE 15° C POINT "A 20° C 5K _XMSN LIMIT 25° C 30° C 35° g POWER FIGURE AC 29.1045-1 ALL-ENGINE CRITICAL ALTITUDE
ADDITIONAL COOLING TEST POINT Page E - 146
AC 29.1047. § 29.1047 (Amendment 29-1) TAKEOFF COOLING TEST PROCEDURES.
a. Objective.
(1) For Category A rotorcraft, the objective is to verify satisfactory takeoff and OEI climb cooling for the Category A takeoff profile defined in aircraft performance §§ 29.59(c) and 29.67(a) following a prolonged hover.
(2) For Category B rotorcraft, the objective is to verify satisfactory cooling for the takeoff and subsequent climb for the Category B takeoff defined in performance §§ 29.63 and 29.65(a) following a prolonged hover.
b. Procedure - Category A.
(1) The rotorcraft is hovered in-ground-effect (IGE) at the power required to hover on the test day at the maximum Category A takeoff gross weight for the hot day, until temperatures stabilize.
(i) Alternate Procedure. If th e test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the measured gas temperature (MGT) indicator. Since operations in the takeoff range are allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and
(ii) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.
(2) After temperatures have stabilized, an OEI climb is initiated from the lowest practicable altitude.
(i) Multiengine power may be used from the stabilized IGE hover to the CDP before OEI operations for cooling verification begin.
(ii) Actual shutdown of the simulated failed engine may not be necessary if the applicant can show that cooling of the remaining engine, fluids, and components is not affected by operation of the “failed” engine at idle power.
(iii) The power utilized at the initiation of the OEI climb would be the same as for establishing the takeoff climbout path of § 29.59, typically 2.5-minute OEI hot-day power available.
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(3) After the time period for which the power is used in establishing the takeoff climbout path has expired, OEI power is changed to that used in meeting the steady rate of climb (150 FPM, 1,000 feet above the takeoff surface of § 29.67(a)(2)).
(i) The power to be used for this phase is 30-minute OEI hot-day power available, if approval of this power rating for performance is requested.
(ii) If 30-minute OEI approval is not requested, the highest hot-day power available approved for continuous usage under OEI conditions would be utilized.
(4) Climb at the OEI power used in meeting § 29.67(a)(2) would continue for at least--
(i) Thirty minutes if 30–minute OEI power is used; or
(ii) Five minutes after the occurrence of the highest temperature recorded, if other than 30-minute OEI is used.
(5) Unlike § 29.1045, the procedure set forth in § 29.1047 for Category A rotorcraft does not specifically require continuation of the OEI climb beyond the 30-minute duration allotted for 30-minute OEI power usage.
c. Procedure - Category B.
(1) The rotorcraft is hovered IGE until temperatures stabilize at the power required on the test day to hover IGE at the maximum Category B takeoff gross weight for the hot day.
(i) Alternate Procedure. If the test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the MGT indicator. Since operation in the takeoff range is allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and
(ii) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.
(2) After temperatures have stabilized in hover IGE, a multiengine climb is initiated at hot-day takeoff power available from the lowest practicable altitude.
Section 29.1047(b)(3) requires only that takeoff power be maintained for the same time interval as used in determining the takeoff flight path under § 29.63. This time interval Page E - 148
could be less than the 5 minutes for which takeoff power is approved. Unless the applicant can show that the time interval used in § 29.63 provides more conservative results, or unless additional testing is proposed, the full 5 minutes allowed for takeoff power should be used to assure that the most critical condition has been surveyed.
(3) After the use of takeoff power for the appropriate time interval, the power should be reduced to multiengine maximum continuous hot-day power available and the climb continued until at least 5 minutes after the occurrence of the highest temperature recorded.
(4) The airspeeds utilized in the climb should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeed most critical for cooling should be used. The airspeed need not exceed the speed for best rate of climb with maximum continuous power available.
AC 29.1049. § 29.1049 HOVERING COOLING TEST PROCEDURES.
a. Objective. The objective is to verify satisfactory hover IGE cooling at sea level and at the hover ceiling for hot-day conditions.
b. Explanation. The rule provides for a hover IGE cooling check in still air at sea level and at the hover ceiling at maximum continuous power. Still air is interpreted as a wind speed of 5 knots or less.
c. Procedures.
(1) The aircraft should be hovered IGE at the maximum certificated hover weight or at the IGE hover weight corresponding to hot-day maximum continuous power available, whichever is less.
(i) The power utilized would normally be hot-day maximum continuous power available and the initial gross weight would be selected as required to achieve hover IGE on the test day.
(ii) After initiation of the hover, special weight control procedures need not be implemented in attempting to maintain hover IGE as fuel burn-off occurs. The power may be gradually reduced to maintain the IGE hover condition.
(2) The hover test is to continue until at least 5 minutes after the occurrence of the highest temperature recorded.
(3) Section 29.1049 also requires a hover IGE at the maximum continuous power available at the altitude resulting in zero rate of climb.
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(i) Often, compliance is illustrated by extrapolating component cooling margins from sea level test results and from selected altitude test site results to the altitude resulting in zero rate of climb.
(ii) Considerable engineering judgment must be exercised in utilizing the extrapolation method described. In general, if test data is extrapolated more than 2,000 feet to the hover ceiling from the highest altitude site selected and the resulting component margin is less than 5° F, additional verification at altitude may be required.
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SUBPART E - POWERPLANT
INDUCTION SYSTEM AC 29.1091. § 29.1091 (Amendment 29-17) AIR INDUCTION.
a. Explanation.
(1) The air induction system for each engine and auxiliary power unit must supply the air required under the operating conditions for which certification is requested. For reciprocating engine installations, the system must provide air that is suitable for proper fuel metering and mixture distribution. This should be shown with the induction system valves in any position.
(2) The intake system shall be designed such that a backfire flame will not constitute a fire hazard within the engine accessory compartment or within other areas of the powerplant compartment.
(3) Each reciprocating engine must have an alternate air source which must be located to prevent entrance of rain, ice, or other foreign matter.
(4) For rotorcraft powered by turbine engines and rotorcraft incorporating auxiliary power units, there must be means to prevent leakage of hazardous amounts of flammable fluids from entering the engine or auxiliary power unit intake system.
(5) Also, the air ducts must be located or protected to minimize the ingestion of foreign matter during takeoff, landing, and taxiing.
b. Procedures.
(1) For turbine-engine installation, the induction system should supply air of suitable quality to meet the installation requirements of the engine manufacturer. The installation requirements should be met throughout the operating envelope of the rotorcraft. In addition, the design and loca tion of the air induction system should prevent accumulations of rain or hail, either external or internal to the induction system, that could adversely affect engine operation.
(2) The inlet design should account for the prevention of hazardous fluids entering the engine. Some designs will have inlet ducts which are free from any fluid lines; however, other designs may route the engine inlet air through a compartment which has flammable fluid lines. When this condition exists, test demonstrations of critical leakage during operation have been used to substantiate the installation. The fluid leakage may not have an adverse effect on engine operation.
(3) The air induction system design should also account for and minimize the possibility of foreign matter ingestion during takeoff, landing, and taxiing.
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AC 29-2C, Chg 4 5/1/2014
(4) For reciprocating engine installations, the induction system should supply air of suitable quality and quantity to the combustion system of the engine. The condition of this air at the entering face of the carburetor is extremely important. For proper operation, it is essential that the airflow be smooth and uniform, clean, and unrestricted throughout the very wide range of horsepower expected from the engine.
AC 29.1093. § 29.1093 (Amendment 29-22) INDUCTION SYSTEM ICING PROTECTION.
a. Reciprocating Engines. No advisory ma terial is presented here for reciprocating engines since it is unlikely that these types will be used in transport rotorcraft.
b. Turbine Engines - Ice Protection.
(1) Explanation.
(i) This rule requires turbine engines and turbine-engine inlets to perform satisfactorily in atmospheric icing conditions defined in Appendix C of part 29. On an equivalent safety basis, the limited icing envelopes described in section 29.1419 of this AC may be used to show compliance with the intent of the regulation if the rotorcraft is limited to not greater than a 10,000-foot pressure altitude for all operations. If operations are permitted above 10,000 feet, the Appendix C, part 29, envelope must be used from 10,000 feet to the service ceiling or 22,000 feet. These possible equivalent safety approaches are not discussed herein. Compliance with the induction system icing protection rule is required regardless of flight manual limitations or restrictions against flight into atmospheric icing conditions.
(ii) In showing compliance with § 29.1093(b)(1)(i), the FAA/AUTHORITY has accepted the concept of limited exposure associated with escape from inadvertent ice encounters.
(A) It is presumed that there will be a flight manual limitation against flight into icing conditions, and that the engine induction system will be reevaluated if total aircraft ice protection certification is requested. Under this concept, the rotorcraft is assumed to fly directly through the icing environment (i.e., direct sequential penetration and straight line exit from both the continuous maximum and intermittent maximum icing clouds). Thus, the duration of exposure to t he icing environment could be calculated by knowing the aircraft flight speed and cloud horizontal extent. A range of engine power and rotorcraft airspeeds should be evaluated to encompass the operating envelope of the rotorcraft.
(B) When this limited exposure concept is used, the aircraft type certificate data sheet should clearly specify that the engine induction system must be reevaluated if certification to the general ice protection regulation, § 29.1419, is requested. This Page E - 152