AC 27-1B, Chg 7 2/4/2016
(iii) Display Integration. The following takes precedence over AC 20-138 (current version) guidance.
(A) Arrangement and visibility.
(1) For rotorcraft installations, section 27.1321 of this AC defines the line of sight and fields of view.
(2) Display of required information (see AC 20-138 (current version) and Society of Automotive Engineer (SAE) Aerospace Recommended Practice (ARP) 4102-7). Like other navigation systems, display the information necessary for the pilot to use the system for its intended function. If providing integration with an autopilot and flight director system, then display the coupled or selected flight director modes in the pilot’s primary field of view. In addition to guidance in AC 20-138, display the following information in the pilot’s primary field of view:
(i) GPS level of service (localized performance with vertical guidance (LPV), lateral navigation/vertical navigation (LNAV/VNAV), LNAV, etc.).
(A) Annunciate other modes not specific to an approach line of minima (i.e., proprietary modes like advisory vertical guidance).
(B) Annunciate mode changes due to degraded GPS capabilities.
(ii) GPS waypoint information.
(A) Waypoint name or identifier.
(B) Distance to and from waypoint.
(iii) GPS alerts and messages (e.g., “SUSPEND,” “MSG”), and any malfunction or downgraded capabilities.
(3) For GPS systems that provide “advisory” vertical guidance, the vertical guidance cueing and symbology should be distinct enough to discriminate it from LPV or ILS vertical guidance. Evaluate the coupling of the autopilot to an advisory glideslope.
(B) Color.
(1) Color philosophy should match the philosophy used on primary flight display and other electronic displays.
(2) For analog displays, colors used in remote annunciation panels should match those used by the GPS manufacturer.
Page MG 1 -14
(iv) Flight Test.
(A) Conduct flight evaluations of the GPS, autopilot, and flight guidance system function together, as required, to ensure integration. Additionally, if use for steep angle point in space (PINS) or other specialized uses are envisioned, flight evaluations are needed to ensure the integration of the activities coupled with the handling characteristics of the aircraft provides the capability to perform the approaches without undue concentration or excessive workload on the part of the pilot. See AC 20-138 (current version), Chapter 20 for further guidance.
(B) Steep angle (> 4 degrees glide path angle) low speed (<70kts KIAS) approaches should be evaluated for IFR certificated aircraft.
(1) Evaluate:
(i) Aircraft handling in different coupled modes at the rotorcraft Vmini at critical CG.
(ii) Steepest glide path angles requested at each of the coupled modes used (see section AC 27, Appendix B of this AC for glide path angle limitations and Vmini).
(iii) For three axis autopilots, assess pilot workload and autopilot capabilities in all coupled modes used for steep approaches.
(iv) For four axis autopilots, assess pilot workload, and autopilot capabilities in all coupled modes.
(v) Assess pilot workload with pilot hand flying:
(A) Minimum equipment required for IFR flight.
(B) Flight director modes, if installed.
(vi) Assess pilot workload, pilot-system interface on missed approaches terminating at the holding fix.
(vii) Abuse testing should include:
(A) Deviation from glidepath and lateral course on intermediate and the final approach segments.
(a) Deviations should be at least one dot laterally and one dot vertically.
Page MG 1 -15
AC 27-1B, Chg 7 2/4/2016
(b) Approach segment flight path and glide path should be recaptured and stabilized prior to the final approach fix (FAF) for the intermediate segment and the Missed Approach Point (MAP) for the final segment.
(c) Vertical speed in the descent should be less than 1,000 fpm.
(d) Torque values should be high enough to keep from entering autorotation.
(e) The pilot should have enough time to easily transition to the visual segment (B) Ability to maintain airspeed at or below the maximum airspeed for the approach.
(C) Maximum allowable tail winds (many steep angle approaches have only one inbound course).
(D) Effect of approach downgraded from LPV to LNAV before and after the FAF, which may require either continuation of the approach or execution of a missed approach.
(viii) Assess the pilot workload and procedures, with autopilot coupled, on complete approaches using transition waypoints and vectors-to-final.
(A) For steepest glide path angle and approach airspeed for which approval is requested.
(B) Assess the handling qualities with the autopilot coupled. If a 4-axis autopilot is proposed, then evaluate the loss of each axis on approach procedures and aircraft handling characteristics.
(C) Assess the transition from coupled instrument flight to uncoupled visual flight at the Decision Altitude (DA).
(a) Accomplish the continuation of the approach in the visual mode to assess the workload required to slow the aircraft at a point to make a landing.
(b) For rotorcraft with Category A procedures, ensure the aircraft can slow to the appropriate speed by the designated landing decision point from the DA.
(ix) If not evaluating the steep angle approaches for IFR aircraft, then apply a limitation in the rotorcraft flight manual or rotorcraft flight manual supplement limiting IFR coupled RNAV approach operations to a maximum glidepath Page MG 1 -16
angle of 3.5° and the minimum approach airspeed that meet flight manual limitations.
This is necessary until evaluations are accomplished and the determination is made that the autopilot-GPS integration supports steep-angle, low speed operations.
(2) Failure mode evaluations:
(i) Evaluate the loss of SBAS and all flight director or vertical guidance display behavior during all phases of operation.
(A) Indications of failure of the SBAS or loss of vertical guidance should be clear and unambiguous to the pilot.
(B) The pilot should be able to transition to degraded approach procedures without undue concentration or exceptional skill.
(ii) Evaluate misleading guidance information. Evaluate performance of advisory vertical guidance provided for LNAV approaches when provided in accordance with AC 20-138 (current version) guidance.
(19) Rotorcraft Condition Monitoring System Installations.
(i) General. There are installations of avionic equipment and systems in rotorcraft to collect data for use in assessing engine and rotorcraft performance, and frequency of maintenance. Some of the items monitored are engine operating exceedances, hot starts, power assurance, and cycle counts. The monitoring systems addressed by this paragraph are those used to collect data for maintenance purposes and not those used in autopilot, flight control, or engine control systems. At present, there are optional approvals requests for most of these systems not performing any required functions. However, most of the applicants anticipate requesting approval for the system use in the future to perform some required function or to allow required maintenance predication 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 b.(19)(iii)(B) below. Refer to section AC 27 MG 15 of this AC for guidance on certification of these systems, also referred to as Health Usage Monitoring Systems (HUMS).
(ii) System Installation. Show that system installations are free from hazards for both normal operation and possible malfunctions. Malfunctions which might be caused by software errors are discussed under paragraph b.(19)(iii)(B) below. The accuracy and response of the monitoring device or system should be sufficient to allow the operational and maintenance personnel to relate the data obtained to required maintenance actions. Use the exceedance (engine limit) information acquired by these systems 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 Page MG 1 -17
AC 27-1B, Chg 7 2/4/2016 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.
(iii) System Components.
(A) Hardware. The system hardware, including airborne electronic hardware, operated 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 RTCA Document DO-160 (current version) should be performed.
(B) Software. If the function of the monitor system depends on embedded airborne software to determine all or part of its functioning, RTCA Document DO-178 (current version) 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 section
27.1309 of this AC. Carefully consider the selection of the software level because sometimes the initial system approval basis is on the system being a non-required optional system. Further, if it shows that there is no dependence on the system software to preclude a hazardous failure mode, then a low software level would be acceptable. However, once initially certified, it is very difficult to qualify software to higher levels of “quality.” Because of this, we recommend choosing the software at the level consistent with the ultimate use for the planned system approval. If the system approval is only to be as non-required optional equipment, then the choice of a low level of software qualification may be appropriate. However, when gaining more experience with the operation of the system, and ultimately planning to seek approval to perform required functions, then initially obtain an appropriate higher level of software.
Note: Consider extensive service experience as a basis for level of criticality without accomplishing RTCA Document DO-178 procedures.
(20) Rotorcraft Health and Usage Monitoring Systems (HUMS).
(i) General. There can be a division of HUMS 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. There are no HUMS approvals to replace service life or other specific physical limits but several systems are now in the process of seeking approval. We consider health monitoring systems the serious applications of this technology, and it will probably be some time before the necessary database 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 Page MG 1 -18
monitoring system described in paragraph b.(19) above. Refer to section AC 27 MG 15 of this AC for guidance on the certification of HUMS.
(ii) Health Monitoring Systems.
(A) We anticipate that 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 level A or equivalent software as defined in RTCA Document DO-178.
(B) Some systems that are under consideration will utilize off aircraft processing of data. If pursuing this, assume 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, specify the intervals for processing the data from each flight as part of the approval.
(C) Due to the limited experience with these systems, we suggest using the issue paper process to record the progress of the approval, and to provide information for later updating of this guidance material.
Page MG 1 -19
CHAPTER 3
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG)
AC 27 MG 2. 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.
(ii) Parallel Load Division.
(iii) 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 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 demonstrat e his equipment either on the bench or installed for ground tests. When a bench setup is Page MG 2 - 1
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 warm-up.
(i) Minimum electrical load for paralleling and minimum engine RPM.
(ii) Vary RPM of all engines from low to high and back to low.
(iii) 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.
(iii) Maximum electrical load and minimum engine RPM.
(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.
Page MG 2 - 2
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).
(i) 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.
(i) 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.
(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 be necessary. Analysis/testing is typically only required for IFR and CAT A complex electrical systems, not basic Part 27 systems. This satisfies the need to verify all analysis, and complex designs are usually verified by tests. IFR and CAT A electrical systems are complex in nature and tests are required to verify the analysis.
Page MG 2 - 3
Additionally, IFR and CAT A requirements refer back to Part 29 requirements for systems and these systems must meet their Part 29.1309 requirements.
(ii) 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.
(iii) 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, 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.
Page MG 2 - 4
(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 AC 27.1309.)
(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.
(ii) 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).
(ii) Instrumentation list (including calibration records).
(iii) Test result recordings.
(iv) Detailed procedures and results obtained.
(v) Conformity inspection records.
(vi) Other data, photographs, etc., to describe the test setup.
(vii) Summary of the test results. This summary should show the maximum load to which each bus, alternator/generator, etc., has been tested.
Page MG 2 - 5
(viii) 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 AC 27.1351.
(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.
(iii) 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.
(ix) Absence of undesired functions in relation to switch combinations.
Page MG 2 - 6