reconstruction of any sequence of events that would happen as a result of any system testing described herein.
(3) Regulatory References. Sections 27.1301, 27.1307(c), (d), (e), 27.1309,
27.1351, 27.1353, 27.1357, 27.1361, 27.1365, and 27.1367.
(4) Miscellaneous, The assigned FAA/AUTHORITY systems and equipment engineer normally witnesses these tests and should be notified as far in advance of the testing as possible to minimize scheduling problems. Conformity of the test setup must be established prior to conducting any testing. Most of the above test categories can be conducted on a bench test setup. A bench test setup is especially recommended in the case of the system malfunction test. It is the applicant's option to demonstrate his equipment either on the bench or installed for ground tests. When a bench setup is used, it should represent the actual aircraft installation to the extent that components and wiring (type, gage, and length) are duplicated. Some retesting may be necessary on the aircraft to verify the bench test results.
b. Ground and Bench Test Procedures.
CAUTION: Prior to disconnecting the battery and removing or adding large loads, either isolate the avionics systems or assure that transients induced are within limits of the avionics equipment.
(1) Normal System Operation.
NOTE: Equipment should be operated for at least 10 minutes prior to each test as a warmup.
(i) Minimum electrical load for paralleling and minimum engine RPM.
(il) Vary RPM of all engines from low to high and back to low.
(ili) Repeat b(1)(ii) for maximum and 50 percent of maximum electrical loads.
(2) Parallel Load Division (if parallel system).
(i) Minimum electrical load for paralleling and minimum engine RPM.
(ii) Fifty percent of maximum electrical load and minimum engine RPM.
(ill) Maximum electrical load and minimum engine RPM.
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(iv) Minimum electrical load for paralleling, vary No. 1 engine RPM from low to high and back to low while holding the RPM of the other engine at minimum (low).
(V) Repeat b(2) (iv) for each other engine on the rotorcraft.
(vi) Repeat b(2) (iv) and b(2)(v) procedures with 50 percent of maximum electrical load.
(vii) Repeat b(2) (iv) and b(2)(v) procedures with a maximum electrical load.
(3) Excitation.
NOTE: All of these tests are to be conducted with the battery OFF since the purpose of the tests is to determine if the ship's battery is necessary for excitation of the alternator(s)/generator(s).
Minimum anticipated electrical load, low engine RPM, and alternator(s)/generator(s) OFF. Demonstrate that when an alternator/generator is turned ON, it will come on the line. Repeat for any other alternators/generators in the system.
(ii) Maximum electrical load, low engine RPM, and alternator(s)/generator(s) OFF. Demonstrate that each alternator/generator will individually come on the line.
(iii) Minimum anticipated electrical load, high engine RPM, and alternator(s)/generator(s) OFF. Demonstrate that each alternator/generator will individually come on the line.
(4) Stabilization.
NOTE: All of these tests are to be conducted with the ship's battery OFF, since the purpose of the tests is to determine if the ship's battery is necessary for stabilization of the alternator/generator. In each case, if the ship's battery is not necessary for stabilization, the alternator/generator should be on the line and remain there at a satisfactory voltage level.
Minimum anticipated electrical load, low engine RPM, alternator(s)/generator(s) ON. Switch on the heaviest electrical load that is anticipated to be installed on the aircraft.
(ii) Repeat b(4)(i) for a maximum electrical load and low engine RPM.
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(iii) Repeat b(4)(i) for a minimum anticipated electrical load and high engine RPM.
(iv) Repeat b(4)(i) for a maximum electrical load and high engine RPM.
(5) System Malfunctions.
(i) System malfunction testing should be conducted as necessary to verify a fault analysis of the proposed electrical system. Depending on the complexity of the proposed system and the comprehensive nature of the system fault analysis, these tests may not be necessary. For Part 27 applications, where less reliance is placed on the functioning of the electrical system for continued safe operation, it is appropriate to rely more heavily on analysis and physical protection and less on testing and technology to account for system malfunctions.
(її) Overcurrent faults (faults to airframe ground that are less than 5.0 Milliohms) should be applied to buses and feeders as necessary to demonstrate that the system's overcurrent circuit protective devices are properly coordinated and provide adequate protection/fault isolation.
Simulate an overvoltage condition on each alternator/generator to demonstrate satisfactory operation of the overvoltage sensing network. On a multiengine configuration, the faulty alternator/generator should be removed without affecting operation of the remainder of the system.
(iv) As appropriate, the annunciation circuitry should be checked for indication of failures such as overvoltage, tripped generators, overcurrent, open feeders, open tie breakers, etc.
(6) Aircraft Ground Tests. If the above tests (reference b(1) through (4) inclusive) are conducted on a bench setup, enough tests should be repeated on the aircraft to validate the bench test results. The following tests should be conducted on the aircraft:
(i) Normal Battery Starts. Start all engines on the aircraft following the normal procedure prescribed in the flight manual. Record starter volts and amperes, time, and any other parameters deemed necessary.
(ii) Ground Power Cart Starts. If the aircraft is equipped with a plug for a ground power cart, use the procedure described in the flight manual and start all engines: Record starter volts and amperes, time, and any other parameters deemed necessary.
(iii) Emergency Battery Operation (if provided). The emergency battery mode of operation should be tested to assure at least proper switching, annunciation, Par 777 1065
and battery capacity. In some instances, an analysis of battery capacity may be adequate.
(iv) Distribution System Tests. With all systems operating individually, open and close feeder circuit breakers and system circuit breakers and assure separation of power sources. If the opening of the feeder protection has been satisfactorily demonstrated on a bench test facility, it should not be necessary to repeat that demonstration on the actual aircraft. The effect of loss of power sources should also be demonstrated on the aircraft.
(v) Other Tests. Conduct other tests as necessary to demonstrate proper operation of the specific design being evaluated.
(7) Environmental Qualification. Each component of the system, such as relays, switches, alternator, generator, sensor, regulator, diode, etc., should be qualified to the critical environmental parameters. The temperature, altitude, humidity, and vibration expected in the approved aircraft operational envelope should fall within those limits the applicant substantiates for the electrical system components. (Refer to Paragraph 621 of this AC.)
(8) Electromagnetic Compatibility. At no time during any of the qualification testing described herein should objectionable interference in the aircraft's radio, navigation, cockpit instrument, autopilot, or interphone system be considered acceptable.
NOTE: The quantitative type testing used for Paragraphs b(7) and (8) above is outside the scope of this document. The latest revision of RTCA Document DO-160 is an acceptable standard.
(9) Transient Tests. The dc system should be tested and shown to exhibit surge, ripple, and spike voltages within the limits of the latest revision of RTCA Document DO-160.
(i) The surge and ripple voltage tolerance of avionic equipment is defined by section 16 of the latest revision of RTCA Document DO-160. Category Z is considered applicable to rotorcraft dc systems.
The voltage spike tolerance of avionic equipment is defined by section 17 of the latest revision of RTCA Document DO-160.
(10) Ground and Bench Test Report. At the conclusion of the ground and bench test program, a report should be prepared and submitted that contains at least the following:
(i) System schematic (including instrumentation tie-in).
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Instrumentation list (including calibration records).
(ilI) Test result recordings.
(iv) Detailed procedures and results obtained.
(v) Conformity inspection records.
(vi) Other data, photographs, etc., to describe the test setup.
(vil) Summary of the test results. This summary should show the maximum load to which each bus, alternator/generator, etc., has been tested.
(viil) Analysis of test results. This should describe how compliance with the regulations has been shown.
c. Flight Test Procedures.
(1) Alternator/generator cooling tests should be conducted in accordance with Paragraph 778 of this AC.
(2) A cockpit evaluation of the electrical system should be conducted to evaluate:
(i) Switch, circuit breaker, and annunciator identification.
(ii) Visibility of placarding, switches, etc., during bright sunlight and night operation.
Color of annunciators as related to the function/malfunction annunciated.
(iv) Load meter readability, if appropriate.
(v) Access to essential switches, circuit breakers, etc.
(vi) Electromagnetic interference.
(vii) Compatibility of the electrical system with the rotorcraft flight manual and the need for additional procedures in the rotorcraft flight manual.
(viii) Clarity of functions such as opened feeder breakers, tie breakers, related annunciation, and necessary corrective action in the event of malfunction.
Par 777 1067
(ix) Absence of undesired functions in relation to switch combinations.
778. STANDARDIZED TEST PROCEDURE FOR ROTORCRAFT GENERATOR COOLING.
a. Test Requirements.
(1) General. The applicant should contact the generator (alternator) manufacturer and obtain the maximum limits for the unit to be tested. This will normally be in terms of temperatures at various locations within the unit (stator, bearings, diodes, heat sinks, brushes, etc.) or in terms of pressure drop across the generator. The manufacturer should either supply an instrumented unit or give complete details for instrumenting the test unit.
(2) Instrumentation Load Bank. A load bank will usually be necessary to load the test unit to the amperage limit for which approval is requested.
(il) Ammeter. An ammeter should be provided with sufficient resolution to assure the amperage load is being maintained at the desired level.
(ili) Temperature/Pressure Readouts. Readouts which are compatible with the temperature or pressure sensors installed in the test unit should be provided.
(iv) Calibration Records. Calibration records should be available for all instrumentation.
(v) Recordings. Permanent recordings should be provided for time, temperatures, current, and/or pressure. The recording device should have provisions for placing event marks on the recording medium.
(3) Regulatory References. Sections 27.1301, 27.1309, 27.1351, 27.1521(f),
27.1041, 27.1043, 27.1045 (through Amendment 27-19).
(4) Miscellaneous. The results obtained from the tests should be corrected for hot-day conditions using a standard lapse rate (3.6° F/1,000 feet).
b. Test Procedures.
(1) Single-Engine Procedure.
The cooling test is to be conducted during ground operation and climb-out, cruise, and approach flight regimes.
1068 Par 777
All ground operational and in-ground-effect hover tests should be conducted in ambient winds of 5 knots or less.
(ill) The battery may be connected to the bus during the generator/alternator cooling test. The generator/alternator temperatures should be recorded at intervals sufficiently close to show the rate of temperature increase and stabilization. The temperature may be considered stabilized when it peaks and has not increased in the last 5 minutes. The climb-out speed and power setting should correspond to the best rate of climb speed, using maximum continuous power or any other normal conditions of climb that would cause the generator/alternator temperatures to be critical. The cruise test should be conducted at maximum altitude in the cruise configuration. Generator/alternator cooling should be conducted at rated output consistent with the RPM at which it is operating. For instance, during the ground tests the engine RPM may be lower than that necessary to sustain maximum rated amperage output. In this case the maximum amperage output of the generator/alternator corresponding to the lower RPM should be assured.
(iv) The test sequence should begin with about 30 minutes of ground operation to account for taxi and holding times and end 5 minutes after all temperatures have peaked after engine shutdown.
(2) Multiengine Procedures. Conduct a generator cooling test in accordance with the following procedures:
All ground operational and in-ground-effect hover tests should be conducted in ambient winds of 5 knots or less.
After engine start, load the instrumented generator to its proposed amperage limit and begin recording temperatures.
(ili) A total of 30 minutes should be spent on the ground prior to takeoff.
This is to account for taxi and holding times.
(iv) After takeoff, climb at single-engine, best-rate-of-climb speed using maximum continuous power, to the single-engine service ceiling. Above this, continue at twin-engine, best-rate-of-climb speed, using maximum continuous power on both engines, to maximum altitude.
(v) Cruise at maximum altitude until all generator temperatures stabilize.
Temperatures shall be considered stabilized when they have peaked and have not increased for a period of 5 minutes.
(vi) Descend, conduct an approach to include a go-around, hover until temperature stabilizes, then land and continue to record temperatures after shutdown until 5 minutes after all temperatures have peaked.
Par 778 1069
(vii) Conduct cooling tests with the rotorcraft hovering at both the minimum and maximum hover altitudes.
(vili) 3 for hot day conditions. Use the standard lapse rate of 3.6° F/1,000 feet for consideration of altitude. See Paragraph 621 of this advisory circular for details on temperature correction.
(ix) If at any time during the testing it appears the manufacturer's limits are to be exceeded, the amperage load on the test generator/alternator should be reduced to prevent this from happening.
779. Sections 27.1301, 27.1309, and 27.1322 ANNUNCIATOR PANELS.
a. Explanation.
(1) The annunciator panel design should be reviewed for the presence of failure modes that can cause illumination of multiple panel segments.
(2) Many test circuits that are diode isolated are vulnerable to this condition. A typical sequence begins with the shorting of a test circuit diode. This failure is undetectable and goes unnoticed until an actual failure condition occurs which causes the associated panel segment to illuminate. At this time all panel segments connected to the test circuit will illuminate.
(3) This configuration becomes a special problem when one or more of the panel segments are red. A red light calls for immediate action by the crew, and the crew does not have adequate information for immediate action when many false panel segments are illuminated.
(4) If the design review indicates a problem, a redesign of the panel to eliminate the condition is considered to be the best solution and is highly encouraged.
b. Procedures
(1) An alternative to panel redesign might be the following:
(• Review the annunciator panel design and note which segments are red.
Determine if cross reference information is available in the cockpit to allow elimination from consideration of any of the red segments (e.g., red low fuel pressure light and low fuel pressure gauge). Normal operation of the gauge would be a reason to assume the light did not cause the problem.
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(III) Where a cross reference is available, further design review of that function is not necessary; however, it may be appropriate to include procedural information in the emergency procedures section of the rotorcraft flight manual.
(iv) If cross references are not available for red segments, additional isolation should be incorporated into the annunciator design for those functions.
(2) If cross referencing is not practical, the following approach is encouraged.
(i) Review the annunciator panel design and note which segments are red.
(ii) Determine if isolation diodes are checked during the application of battery or external power before starting the engines (e.g., red low oil pressure light). If isolation diode is shorted, all panel segments will light as soon as battery or external power is applied.
(iii) When the isolation diode can be checked before starting engines, further design review is not necessary.
(iv) If diodes are not automatically checked before starting, then additional isolation, should be considered.
c. Annunciator Panel Arrangement. The annunciator panels should be arranged in a logical manner to reduce the crew's time required to locate faults and to increase their efficiency in following aircraft flight manual procedures. For example, engine annunciators on multiengine rotorcraft should be physically located on the panel to coincide with engine location (left or right) so that properly operating engines are not inadvertently shut down due to crew confusion over which engine has malfunctioned.
Par 779 1071
АС 27-1A 7/30/97
780. DEFINITION OF ENGINE ISOLATION FEATURES AS APPLIED TO §§ 27.79(b)(2), 27.141(b)(1), and 27.143(d)(1) (Amendment 27-19).
a. Explanation.
(1) Each of the cited performance and flight characteristic sections of Part 27 mention multiengine rotorcraft meeting transport Category A engine isolation requirements or refer to engine isolation features which ensure continued operation of the remaining engine. Unlike normal category fixed-wing (Part 23, § 23.903(c)) and the transport category fixed-wind and rotorcraft regulations, Part 27 does not provide a general engine isolation rule to make this determination.
(2) While it is clear that Part 27 does not require complete engine isolation, if credit for this feature is claimed (i.e., sudden complete engine power failure is not considered in showing compliance with the cited section), criteria must be established to allow a satisfactory isolation assessment.
(3) An approach which the FAA/AUTHORITY would find acceptable in making a Part 27 engine isolation determination is given. The FAA/AUTHORITY logic for establishing this criteria is also presented.
b. Criteria.
(1) The engine isolation provided may be for an appropriate limited time period.
(2) The failure or malfunction of any engine or the failure of any system that can affect any engine will not--
(i) Prevent the continued safe operation of the remaining engines for the appropriate limited time period; or
(ii) Require immediate action by a crewmember for continued safe operation.
(3) Each engine must be isolated by a firewall, shroud, or equivalent means from the remaining engines.
c. Criteria Rationale.
(1) Category A Minimum Time for Isolation. The acceptance of a limited time period for engine isolation is consistent with the acceptance of a reduced level of safety for a Part 27 rotorcraft. The criteria is also consistent with the Part 27 philosophy of allowing for a controlled landing following engine failure versus the Part 29 Category A principle of continued safe flight and a controlled landing.
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(2) Installation Analysis.
(i> The degree of engine isolation can be established by an installation assessment against the § 29.903(b) general isolation requirement, as modified for Part 27 by a limited time period concept.
Table 780-1 is a listing of the Part 29 sections that may be involved in Category A engine isolation considerations. Sections 29.901(c) and 29.903(b) are the general isolation regulations under which the other more specific rules naturally fall.
The point that the selection of specific rules from Table 780-1 does not achieve the desired degree of transport Category A engine isolation, and that the general isolation rules (89 29.901(c) and 29.903(b)) must be used, is illustrated by the following examples.
(A) Example #1. No specific requirement from Part 29 (or Table 780-1) can be cited which precludes a common engine mount. The design of the mount could be such that its failure results in sudden, complete power loss from all engines.
(B) Example #2. No specific Part 29 requirement prohibits a common engine induction system. F.O.D., fire in the induction system, or the adverse affect of engine surge on the remaining engine could result in sudden, complete power loss from all engines.
(C) Example #3. Crosstalk between engine fuel controls (possibly used for power matching) or the use of a common input parameter signal to the fuel controls is not prohibited by any specific isolation rule. Signals could be received which command the simultaneous shutdown of all engines.
(iii) These examples clearly illustrate that specific Part 29 isolation rules cannot be selected to establish appropriate Part 27 engine isolation, and that the installation must be evaluated by the general isolation policy set forth. This can be readily accomplished by a failure mode and effects analysis (FMEA).
(3) Firewalls.
CAR 6.483, prior to Amendment 6-4, effective May 15, 1953, requires "All engines, auxiliary power units, fuel burning heaters, and other combustion equipment which are intended for operation in flight shall be isolated from the remainder of the rotorcraft by means of firewalls, shrouds, or other equivalent means."
This rule would clearly require a firewall between engines of multiengine rotorcraft.
(ii) Amendment 6-4 revised § 6.483 to read "Engines shall be isolated from personnel compartments by means of firewalls, shrouds, or other equivalent means. They shall be similarly isolated from the structure, controls, rotor mechanism, Par 780 1073
АС 27-1А 7/30/97 and other parts essential to a controlled landing of the rotorcraft..." (remainder essentially identical to current § 27.1191).
(ії) The preamble explanation of Amendment 6-4 states that these changes are "intended to afford greater protection to the crew and passengers in the event of fire during flighti This revision did not intend to authorize less firewall isolation between the engines than was required by the earlier version. Also, the subsequent paragraphs clearly require firewalls between other combustion equipment and the rest of the rotorcraft (§ 27.1191(b)). To accept anything less for the engine is clearly inappropriate. Further, § 23.1191 requires firewalls or equivalent means between each engine and the rest of the airplane, and current safety requirements pertaining to in-flight fires should be no less stringent for normal category rotorcraft.
(iv) The lack of a firewall between engines or any other design arrangements which, in the event of one engine failure creates definable jeopardy for the remaining engines, will result in a significantly lower level of safety than is being assumed by the operators.
(v) A regulation change to clarity this § 27.1191 rule is planned.
d. Guidance.
(1) The minimum appropriate limited time period of engine isolation which would allow establishment of a one-engine-inoperative HV diagram, § 27.79(b)(2), would be defined by the time increment to recognize the engine failure and to make a landing from the most critical point on the desired HV diagram.
(2) The minimum appropriate limited time period of engine isolation to show compliance with §§ 27.141(b)(1) and 27.143(d)(1) considering the sudden power failure of one engine (rather than sudden complete power failure) would be the time increment to recognize the engine failure and to transition to a flight condition where failure of the remaining engine can be tolerated.
(3) Some existing provisions of Part 27 require isolation of certain systems (oil, fuel, and engine controls) without regard to a limited time period. These existing Part 27 engine isolation provisions must be observed regardless of the policy discussed herein.
(4) The limited time period concept must not be utilized to eliminate protection otherwise required by specific rules of Part 27 or to reduce accepted test conditions.
For example, lines which carry flammable fluids in areas subject to engine fire conditions must be fire resistant (§ 27.1183(a)). Fire resistant hose standards require testing for at least 5 minutes at 2000° F.
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