(B) Communicate with the ground radar personnel for evidence of transponder dropout. During the flight, check the "ident" mode of the ATC transponder to assure that it is performing its intended function. Determine that the transponder system does not interfere with other systems (except possibly the DME) aboard the rotorcraft and that other equipment (except possibly the DME) does not interfere with the operation of the transponder system. There should be no dropouts, that is, when there is no return for two or more sweeps. The operation of the ATC transponder should be verified over the station at 25 NM and at 80 NM.
(C) Long Range Reception. Perform two 360° turns, one to the right and one to the left, at bank angles of 8° to 10° with the flight pattern 80 NM from the radar facility. During these turns, the radar display should be monitored and there should be no signal dropouts (two or more sweeps).
(10) Weather Radar Equipment.
Bearing Accuracy. The indicated bearing of objects shown on the display should be within 5° of their actual magnetic bearing within the sectors 40° right and left of the aircraft longitudinal axis. Beyond 40° right and left, bearing accuracy should be 110°.
(il) Distance of Operation. The radar should be capable of displaying prominent targets throughout the distance and angular range of the display.
(¡1) Antenna Stabilization. When antenna stabilization is provided, it should eliminate blurring of the display for the ranges of pitch and roll for which it is designed.
(iv) Beam Tilting. The radar antenna should be installed so that its beam is adjustable to any position between 10° above and 10° below the plane of rotation of the antenna.
(v) Technique.
(A) Bearing Accuracy. Fly under conditions which allow visual identification of a target, such as an island, a river, or a lake, at a range within 10 percent of the maximum range of the radar. When flying toward the target, select a course that will pass over a reference point from which the bearing to the target is known. When flying a course from the reference point to the target, determine the error in displayed bearing to the target on all range settings. Change heading in increments of 10° and determine the error in the displayed bearing to the target.
(B) Contour Display (so Echo). If heavy cloud formations or rainstorms are reported within a reasonable distance from the test base, select the contour display mode. The radar should differentiate between heavy and light 1056 Par 776
7/30/97 АС 27-1A precipitation. In the absence of the above weather conditions, determine the effectiveness of the contour display function by switching from normal to contour display while observing large objects of varying brightness on the indicator. The brightest objects should become the darkest when switching from normal to contour mode.
(C) Stability. While observing a target return on the radar indicator, turn off the stabilizing function and put the aircraft through pitch and roll movements.
Observe the blurring of the display. Turn the stabilizing mechanism on and repeat the roll and pitch movements. Evaluate the effectiveness of the stabilizing function in maintaining a sharp display.
(D) Ground Mapping. Fly over areas containing large, easily identifiable landmarks such as rivers, towns, islands, coastlines, etc. Compare the form of these objects on the indicator with their actual shape as visually observed from the cockpit.
(E) Mutual Interference. Determine that no objectionable interference is present on the radar indicator from any electrical or radio/navigational equipment when operating, and that the radar installation does not interfere with the operation of any of the rotorcraft's radio/navigational systems.
(11) Area Navigation. Advisory Circular 90-45A is the basic criteria for evaluating an area navigation system, including acceptable means of compliance to the FAR.
(12) Inertial Navigation. Advisory Circular 25-4 is the basic criteria for the engineering evaluation of an inertial navigation system (INS) and offers acceptable means of compliance with the applicable FAR which contain mandatory requirements in an objective form. The engineering evaluation of an INS should also include awareness of AC 121-13 which presents criteria to be met before an applicant can get operational approval. For flights up to 10 hours, the radial error should not exceed 2 NM per hour of operation on a 95 percent statistical basis. For flights longer than 10 hours, the error should not exceed ‡20 NM cross-track or ‡25 NM along track error.
A 2 NM radial error is represented by a circle, having a radius of 2 NM, centered on the selected destination point.
(13) Doppler Navigation. Doppler Navigation System installed performance should be evaluated in accordance with AC 121-13, Self-Contained Navigation Systems (Long Range). (See Part 121, Appendix G).
(14) Radio Altimeters. Radio Altimeter System installed performance should be evaluated in accordance with RTCA Document DO-123, Appendix A, Part II.
(15) Emergency Locator Transmitters (ELI).
Par 776 1057
(i) ELT performance should be evaluated in accordance with TSO-C91.
ELT installations should be examined for potential operational problems. There have been numerous instances of interaction between ELT and other VHF installations.
ELT antenna installations in close proximity to other VHF antennas should be suspect.
Antenna patterns of previously installed VHF antennas should be measured after an ELT installation. Some problems caused by ELT installations are as follows:
(A) Loss of radiated power from VHF communications.
(B) Reradiation of VHF transmitter energy such that navigation crosspointers are affected.
(C) Reception of FM broadcast, at high level, in VHF communications.
(D) Inadvertent activation of the ELT by VHF transmitted energy.
(See AD 72-22-3.
ELT installation. TSO-C91 specifies that the ELT be automatically activated when subjected to a force of 5.0(+2,-0)g in the direction of the longitudinal axis of the aircraft. This recommendation for mounting is considered satisfactory for rotorcraft. In recognition of the significant vertical impact velocity that rotorcraft commonly have, an optional placement of the ELT pitched down 30° from the horizontal axis of the rotorcraft is also satisfactory.
(16) Audio Interphone Systems. Acceptable communications should be demonstrated for all audio equipment including microphones, speakers, headsets, and interphone amplifiers. All modes of operation should be tested, including operation during emergency conditions (i.e., emergency descent, and oxygen masks) with all rotorcraft engines running, all rotorcraft pulse equipment transmitting, and all electrical equipment operating.
(17) Portable Battery Powered Megaphones (AC 121-6). Megaphone performance should be evaluated in accordance with AC 121-6.
(18) Omega and Omega/LF Navigation Systems. Omega and Omega/LF Navigation systems should be evaluated in accordance with the following advisory circular that applies to the type of approval requested:
AC 120-37, Approval of Omega Systems as a Sole Means of Overwater Long Range Navigation.
(ii) AC 120-31A, Approval of Airborne Omega Navigation Systems as a Means of Updating Self-contained Navigation Systems.
1058 Par 776
(i¡i) AC 20-101B, Approval of Omega and Omega/VLF Navigation.
(19) Rotorcraft Condition Monitoring System Installations.
General. Avionic equipment and systems are being installed in rotorcraft to collect data to be used in assessing engine/rotorcraft performance and frequency of maintenance. Some of the items monitored are engine operating exceedances, hot starts, power assurance, and cycle counts.
(A) The monitoring systems being addressed by this paragraph are those used to collect data for maintenance purposes not those monitors which are utilized as part of the control systems for autopilot/flight controls or engine controls.
(B) At present, optional approvals are being requested for most of these systems not performing any required functions. However, most of the applicants anticipate requesting approval for the system to be used in the future to perform some required function or to allow required maintenance to be predicated on the operation of the system. This consideration becomes particularly important if the system is software based. A further discussion of system software is included in Paragraph 776b(19) (iii)(B).
(ii) System Installation. The system installation should be shown to be free from hazards considering both normal operation and possible malfunctions.
Malfunctions which might be caused by software errors are discussed under Paragraph 776b(19) (ili)(B). The accuracy and response of the monitoring device/system should be sufficient to allow the operational and maintenance personnel to relate the data obtained to required maintenance actions. The exceedance (engine limit) information being acquired by these systems is or will be used in place of information previously acquired from field reports of operational personnel utilizing the basic aircraft instruments. In this case, the automated system will generally produce results which are more accurate than the basic aircraft instruments. However, in this circumstance, it is not appropriate to require the monitor system to be more accurate than the previously approved methods used to provide the required exceedance data.
If the data collected by the system require filtering prior to use, it is equally acceptable to accomplish this filtering either as the data are being acquired (airborne function) or when the data are analyzed (ground based function) and used in the maintenance of the rotorcraft.
(i) System Components.
(A) Hardware. The hardware of the system when operating under the control of the imbedded software should be shown to comply with § 27.1301.
Additionally, in showing compliance to § 27.1309(a), laboratory testing to the appropriate portions of the latest revision of RTCA Document DO-160 should be performed.
Par 776 1059
(B) Software. If the function of the monitor system depends on embedded airborne software to determine all or part of its functioning, Document DO-178 is the recommended standard to be used for the approval of the system software. A further discussion of the use of this document is included in Paragraph 621. The selection of the software level should be carefully considered because system approval is sometimes initially sought on the basis of the system being a non-required optional system. If it has further been shown that no dependence is made on the system software to preclude a hazardous failure mode, then a low software level would be acceptable. However, it is very difficult to qualify software to higher levels of "quality" once the software has been initially certified. Because of this, it is recommended that the software be chosen to the level consistent with the ultimate use to which approval of the system is planned. If the system is to be approved only as non-required optional equipment, then the choice of a low level of software qualification may be appropriate. However, when more experience is gained with the operation of the system, and it is ultimately planned to seek approval to perform required functions, then an appropriate higher level of software should be initially obtained.
NOTE: Extensive service experience should not be considered as a basis for level of criticality without accomplishing RTCA DO-178 procedures.
(20) Night Vision Goggles (NVG).
Background. Night vision goggles (NVG) have been used by U.S.
military pilots since the early 1970's. The first units (first generation or GEN I) were constructed from the rifle "Sniper-Scopes." These units did not provide much light amplification. The second generation (GEN II) were still primarily designed for ground use. Second generation high performance units (military designation AN/PVS-5C) had some consideration for flight use but were still lacking in several aspects. A light level of at least a quarter moon well above the horizon was required for operation of these NVG. At first the normally helmet-mounted units covered the pilots entire upper face and the pilot could only see through the NVG. In order to protect the light amplification system these NVG had an automatic shutoff feature when brighter than relatively low levels of light were encountered. Normal incandescent and especially red incandescent lights would cause these NVG to shut down. Aircraft cockpit lights, especially the red warning lights, would cause "blooming" (an increased brightness of all or portions of the NG field of view with the disappearance of the "picture" in that area or a total shutdown of the NVG. Military aircraft cockpits and lighting systems were significantly modified to avoid this problem: In the late 1980's the military pushed technology for better and aircraft compatible NVG. Third generation (GEN III, military designation ANVIS or AN/AVS-6) NVG systems became available about 1988. These systems require only star light for satisfactory operation.
(i) Procedures. As of January 1990, no approvals for civil rotorcraft operations with NVG have been issued. Since NG are not installed in the rotorcraft, 1060 Par 776
7/30/97 АС 27-1А they are not required to be approved as part of the type design. However, since an operational approval would be required for use of NVG, they should meet some acceptable performance standard. The minimum standard recommended is the GEN III NG. The performance of these NG are rated as their spectral response to irradiated light sources, measured as density of incident photons per square meter. Third generation, AN/AVS-6, NVG have been evaluated for compatibility with a limited number of rotorcraft and were generally found to be usable during en route operations with no cockpit lighting systems modifications. It is anticipated, however, that some aircraft may require significant modification to the existing cockpit lighting systems. The FAA/AUTHORITY policy is that modification of the cockpit to a non-compliant configuration to accommodate NVG use is not acceptable. For instance, alteration of the required red warning annunciators to some other color is not acceptable. Since individual rotorcraft may have been modified with additional lights or systems, each rotorcraft being considered for use with NVG should be evaluated by an FAA/AUTHORITY representative during a night flight. If it is anticipated that cockpit lighting system modifications will be required to achieve an adequate level of NVG compatibility FAA/AUTHORITY involvement should be arranged as soon as possible.
Preferably this evaluation flight would be made with two pilots or a pilot and safety observer, over a known area, where all the aircraft and cockpit lights are operated and their effect on the NVG determined. Reflections of landing or searchlights on windshields or other glass during approach or landing may affect NVG and may impose a minimum altitude restriction for use of NVG. Failure of the NVG should be evaluated during any critical flight phase.
Note that the above discussion is purposely limited in scope. Issues such as crew training and operating limitations would have to be addressed in detail to obtain an operational approval.
(21) Rotorcraft Health and Usage Monitoring Systems (HUMS).
General. HUMS can be divided into two major categories: Health Monitoring Systems and Usage Monitoring Systems. The provisions of § 27.1301 are used to determine that the system performs its intended function. The provisions of § 27.1309(a) and (b) are used to look at the impact of environmental conditions and malfunctions. To date (mid-1990) HUMS have not been approved to replace service life or other specific physical limits but several systems are now in the process of seeking approval. Health monitoring systems are considered to be the serious applications of this technology, and it will probably be some time before the necessary data base to allow full reliance on this technology is available. There have been numerous approvals of usage monitoring systems as optional equipment, and a good example of this technology is a condition monitoring system described in 776b(19) above.
(ії) Health Monitoring Systems.
Par 776 1061
(A) It is anticipated these systems will begin as optional" systems in order to build a data base to support expansion of the approval to achieve credit for extension of maintenance intervals, and so forth. Some of these applications may require system redundancy, and some may require DO-178A Level I or equivalent software.
(B) Some systems that are being considered will utilize off aircraft processing of data. If this is to be pursued it should be assumed that the aircraft data will be lost or misplaced at the processing center, and the aircraft system design should consider this possibility. Some on board data storage is one way to account for this lost data. The integrity of the processing center's software should be equal to that of the aircraft software. In addition the intervals for processing the data from each flight should be specified as part of the approval.
(C) Due to the limited experience with these systems it is suggested the issue paper process be utilized to record the progress of the approval, and to provide information for later updating of this AC material.
777. STANDARDIZED TEST PROCEDURE FOR ROTORCRAFT DC ELECTRICAL SYSTEM TESTS.
a. Test Requirements.
(1) General. The following functions and characteristics are to be evaluated:
(i) Normal System Operation.
(il) Parallel Load Division.
(i) Excitation.
(iv) Stabilization.
(v) Systems Malfunction.
(vi) Environmental Capability.
(vii) Electromagnetic Compatibility.
(viii) Cooling Capability.
(ix) Surge Characteristics, Ripple Voltage, and Voltage Spikes.
(2) Instrumentation. Calibration records should be available for all instrumentation. Enough specific currents and voltages should be recorded to allow 1062 Par 776
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.
Par 777 1063
(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.
1064 Par 777
(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).
1066 Par 777
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