(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
to measure the 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 COp 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, througn 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 wind 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.
(Ili) Ambient temperatures.
(iv) Pressure altitude.
(v) Fuel used (or remaining).
(vi) Load cell reading.
(vii) Generator(s) load.
82 Par 56
7/30/97/ AC 29-2B As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in Paragraph b(4) below will determine if any problems, such as load cell malfunctions have occurred. The free flight hover data must fall within the allowable scatter of the tethered data.
(4) If there are no provisions or equipment to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of emergency. The rotorcraft must be reballasted to different weights to allow the maximum COp spread. When using the free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system which may influence hover performance, it must be accounted for.
(5) It is extremely difficult to determine when a rotorcraft is hovering OGE at high altitudes above ground level since there is no ground reference. In a true hover, the rotorcraft will drift with the wind. Numerous techniques have been tried to allow OGE hover data acquisition at high altitudes, all of which have resulted in much data scatter. Until a method is proposed and found acceptable to the FAA/AUTHORITY, OGE hover data must be obtained at the various altitude sites where IGE hover data is obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.
57. § 29.51 TAKEOFF DATA - GENERAL.
a. Explanation. Section 29.51 details the conditions under which takeoff performance data can be obtained and presented in the FAA/AUTHORITY approved flight manual. The flight manual must also contain the technique(s) to be used to obtain the published flight manual takeoff performance. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 29.51. Rotorcraft are different from one another and due to this, 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 and used to achieve the performance 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.
b. Background.
(1) Certain special takeoff techniques are necessary when a rotorcraft is unable to takeoff vertically because of altitude, weight, power effects, or operational Par 56 83
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, altitude, and temperature for which certification is requested.
(2) The primary factor which determines the rotorcraft's takeoff performance is the amount of excess power available. Excess power available is the difference between the power required to hover at the reference height above the ground and the takeoff power available from a minimum installed specification engine. Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV constraints. In such situations, hover power required plus some power increment may be the maximum that can be used and the resulting performance determined accordingly.
(3) Landing gear height above the ground should not be greater than that demonstrated satisfactorily for HV, rejected takeoff, and that height for which IGE hover performance data is presented in the RFM, or less than that height below which ground contact may occur when accomplishing takeoff procedures. For rotorcraft fitted with wheels, a running takeoff procedure may be accepted. The hover reference height is established as the minimum landing gear height above the takeoff surface, from which a takeoff can be accomplished consistently in zero wind without contacting the runway.
Category B takeoff must be accomplished with power fixed at the power required to hover at the reference height (not greater than the height for which IGE performance data is presented).
c. Procedure. There are different techniques which may be used in order to 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 schedule airspeed(s) and constant rotor RPM.
Power adjustments may be accomplished to maintain targeted power except where procedure requires high workload outside 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).
57A. § 29.51 (Amendment 29-39) TAKEOFF DATA - GENERAL.
a. Explanation. Amendment 39 added takeoff requirements in new §§ 29.55,
29.60, 29.61 and 29.62.
84 Par 57