AC 27-1A 7130/97 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 58-6 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 pretakeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated power values. This occurs when the curve of corrected power and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished.
(6) Deteriorated Engine Power - Turboshaft Engine.
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 deteriorated engine power. These include:
(1) Engine characteristics (§ 27.939). Surge margin, engine response, and air-restart capability might be affected and should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism.
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(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 producer speed, and thus power available under certain ambients, may be affected if Pc is an input to the fuel scheduling mechanism.
(4) Systems for surge protection which schedule on Pc such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The effect would normally be negligible unless when installed, the installation losses, combined with reduced Pc because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.
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 (5) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines is greatly diminished and perhaps eliminated.
(B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in Paragraph (5) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur with cold temperatures and high altitudes; bleed valves and other devices which schedule on gas producer discharge pressure are most likely to function and reduce power on a hot day at low altitude.
(C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated engine monitoring procedures will be introduced when allowing the use of deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in-flight procedure must be evaluated by flight tests as described in Paragraph (5) (engine power-turboshaft engines) herein. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure Par 58 69
already in place for the specification engine, which could eliminate the need for flight test evaluation.
58A. § 27.45 (Amendment 27-21) GENERAL.
a. Explanation. Amendment 27-21 adds § 27.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered rotorcraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the 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.
58B. § 27.45 (Amendment 27-29) GENERAL.
a. Explanation. Amendment 27-29 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 Appendix 1 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved.
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CERTIFICATION BASIS Rgmts 27-Amdt. 21 27-Amdt. 2 FAR 27 Original 6-Amdt. 7 6-Amdt. 4 CAR 6 Original Original CAR 06 7/30 / 97
1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. No specific H-V Level Level Level Level Level Level wt. and alt.
00 Ref. 2. Max. OGE 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected requirements.
27.25 wt. Lesser of: by applicant at by applicant at by applicant by applicant by applicant at
27.1519 27.79 Test b. 7000' Hd a. Max. alt. cap. lessor of: a. Max. alt. cap. lessor of: a. Max. alt. cap. a. Max. alt. cap. at lessor of: Idg. alt. max. t.o. and Idg. alt. max. t.o. and
27.1587 Conditions b. 7000' Hd b. 7000' Hd b. 7000' Hd
6.116
6.741
6.743
3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is H-V is Info. info. info. info. info. operating operating
4. Max. alt. For 4. If H-V wt is limitation. limitation.
REM which H-V is valid less than IGE Wt., wt. becomes limitation for
t.o.Adg.
AC 27-1 5. If H-V is less Para than OGE wt., 58 & 59 H-V wt. becomes limit.
6. Applicant is encouraged to demo H-V to Remarks WAT limits.
7. Hover data may be shown above 7000' if AC 27-1A H-V & IGE are demo'd to 7000' FIGURE 58-1. H-V REQUIREMENTS
CERTIFICATION BASIS FAR 27 CAR 06 Ramts 27-Amdt. 21 Original Original Original
1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level 1. No specific CONTROL 27.1587 27.143 27.25 IGE Ref. Conditions Test b. 7000' Hd 2. Max. IGE wt. Lesser of: a. Max. alt. cap. 3. Critical CG 4. Critical Rotor t.o. and Idg. alt. 2. Wt. selected by 3. Critical CG applicant to max. 4. Critical Rotor RPM 2. Max. approved wt. for t.o./dg at alt 3. Critical CG 4. Critical Rotor RPM 5. Wind not less than above sea level. * get ayed in requirement. requirement.
AC 27-1 RPM 5. Critical wt. 20 mph.
Para 5. Wind of not less 6. Wind of not less 58 & 81 than 17 kts. than 20 mph REM 6. Max. safe wind is 7. Max. safe wind 6. Max. safe wind is perf. info. is perf. info. perf. info.
7. If 17 kts. wind demo'd to alt. less Remarks than 7000', a corresponding WAT limit must be established.
7/30/97 FIGURE 58-2. IGE CONTROLLABILITY REQUIREMENTS
600 CORRECTED SHAFT HORSEPOWER 500 ENG S/N 12345- 400 300 MIN CERTIFIED
SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE 58-3. SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE -
SEA LEVEL STANDARD DAY Par 58 73
PRESSURE ALTITUDE - FEET (X 1000) 2 200 300 400 500
SHAFT HORSEPOWER AVAILABLE FIGURE 58-4. UNINSTALLED TAKEOFF POWER AVAILABLE 74 Par 58
PRESSURE ALTITUDE - FEET (× 1000) 7/30/97 Par 58 FIGURE 58-5. INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM 200 50 ° C 40 ° C SHAFT HORSEPOWER AVAILABLE 300 20 ° C 30 ° C 10 ° C 0°C + -10° € -20 ° C 400 -30 ° 10-50 °C TRANSMISSION AC 27-1A LIMIT 75
CONDITIONS: EXAMPLE: AC 27-1A ZERO AIRSPEED OAT = 17.5°C RPM 100% TOT = 760° C ANTI-ICE OFF Hp = 4000 FT GENERATOR LOAD 100 AMPS REQUIRED TORQUE = 70% 1000 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 58 OUTSIDE AIR TEMPERATURE - °C 10 20 30 40 60 20 PERCENT TORQUE 40 60 80 100 120 7/30/97 FIGURE 58-6. POWER ASSURANCE CHECK CHART
59. § 27.51 TAKEOFF.
a. Explanation. Section 27.51 details the conditions under which takeoff data must be obtained. The flight manual must contain the technique(s) to be used to obtain the published flight manual takeoff procedures. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 27.51. Because rotorcraft differ, different pilot techniques are sometimes required to achieve the safest and most optimum takeoff performance. The recommended technique that is published in the flight manual must be determined to be one that the operational pilot can duplicate using the minimum amount of type design cockpit instrumentation and the minimum crew. Only rotorcraft takeoff techniques will be covered in this section.
b. Background.
(1) Certain special takeoff techniques are necessary when a rotorcraft is unable to take off vertically because of altitude, weight, power effects, or operational limitations. The recommended technique used to take off under such conditions is to accelerate the rotorcraft in-ground-effect (IGE) to a predetermined airspeed prior to climbout. Takeoff tests are performed to determine the best repeatable technique(s) for a particular rotorcraft over the range of weight and altitude for which certification is requested.
(2) Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV or aircraft attitude constraints. In such situations, hover power required plus some power increment may be the maximum recommended for use.
(3) Wheel or skid height should be not less than that demonstrated satisfactorily for the high speed, low altitude portion of the HV curve, or that height below which ground contact may occur when accomplishing takeoff procedures.
(4) For rotorcraft fitted with wheels, a running takeoff procedure may be accepted.
c. Procedures.
(1) There are different takeoff profiles which may be used to complete a maximum performance takeoff in a rotorcraft. The manufacturer will normally determine which method is best for a particular rotorcraft. The most commonly accepted method is the hover and level acceleration technique. In this technique, the rotorcraft is stabilized in a hover at the reference height. From the stabilized hover, the rotorcraft is accelerated to the climbout airspeed using the predetermined takeoff power. When the desired climbout airspeed is achieved, the rotorcraft is rotated and the climbout is accomplished at the scheduled airspeed(s) and constant rotor RPM. Power adjustments may be accomplished to maintain the targeted power except where Par 59 77
procedure requires high workload outside the cockpit (i.e., that portion of takeoff where horizontal acceleration close to the ground has pilot scan outside the cockpit and adjustment of engine torque or temperature would require an undue increase in workload). The recommended takeoff procedure must be demonstrated to remain clear of the HV "avoid" areas without requiring exceptional piloting skill or exceptionally favorable conditions.
(2) The hover reference height is established as the minimum skid or wheel height above the takeoff surface from which a takeoff can consistently be accomplished in zero wind without contacting the runway surface. The takeoff must be accomplished with power fixed at the power required to hover at the hover reference height and must not require exceptional piloting skill to avoid runway surface contact.
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60.-62. RESERVED.
Intentionally Left Blank Par60 79 (thru 81)