AC 27.1335. § 27.1335 (Amendment 27-13) FLIGHT DIRECTOR SYSTEMS.
a. Explanation. This section prescribes the accepted display criteria for a rotorcraft three-cue flight director providing command guidance for pitch, roll, and power. Three-cue flight directors for rotorcraft use the usual pitch and roll command cues with the third cue displayed on the left side of the attitude director indicator (ADI).
These instruments can be used in either the two-axes or three-axes modes. In either mode, the lateral command cue controls the roll attitude, and the vertical command cue controls the pitch attitude. The rotorcraft attitude, controlled by the cyclic control, is changed to satisfy the flight director commands. The third cue, when displayed, commands collective pitch position and is used when an airspeed or pitch attitude mode and a vertical mode (altitude hold, glide slope, etc.) are selected.
(1) The general convention for flight director design is that each command bar is a "fly to" command. The motion of the flight director indicator is such to command a corresponding sense of control system motion. This is true of flight director pitch and roll commands and should hold true for additional commands such as collective pitch.
(2) Some consideration should be given to the collective, or third cue, display.
For example, if the collective symbol is selected as the fixed index, the command cue and collective pitch control should move in opposite directions when collective pitch changes are made. This configuration would constitute a conventional "fly to" indicator.
If the collective symbol is selected for the movable index, the direction of motion of the collective symbol will coincide with the direction of collective pitch changes. In this case the moving collective symbol does not comply with the "fly to" convention; however, this configuration has been approved by the FAA with special symbology, special background effects, and special color coding and has performed satisfactorily in service.
b. Procedures. The recommended display for a three-cue flight director incorporates the standard pitch and roll command symbols, either pitch and roll bars or the "V" bar display. The third cue, or collective symbol, should be located on the left side of the ADI. The shape of the moving cue and the background display should be unique to avoid being confused with a glide slope display or angle of attack display.
One display uses a third cue, shaped like a small handle, to aid in identifying it as the collective pitch symbol.
(1) The color of the pitch and roll command indicators, the aircraft symbol, the background marking of the third cue, and third cue itself, should be consistent. The optimum color scheme uses the same color for the aircraft symbol and the collective symbol. This is usually fire orange. The command cues including the collective cue also should use the same color, usually yellow. The rationale for the different colors is that the aircraft symbol and the collective symbol (the same color) are moved toward their respective command cues. If the pitch command cue is above the center, the aircraft symbol is raised (nose pulled up) and, if the collective command cue is above the collective symbol, the collective pitch is raised, moving the collective symbol towards the command cue.
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AC 27-1B, Chg 4 3/31/2014
(2) If the attitude director indicator (ADI) provides a monochromatic display, the collective pitch cue and its background markings must be distinctive to reduce the chance of being confused with the glide slope indicator. This can be accomplished through the use of different shaped cues and background marks. A round cue with a chevron-shaped background marking has been satisfactory.
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AC 27.1335A. § 27.1335 (AMENDMENT 27-51) FLIGHT DIRECTOR SYSTEMS.
Explanation. Modern designs include both automatic pilot and flight director systems, referring to them now as automatic pilot and flight guidance systems. Having Dovisions hersection 27. 1ament 27-51 removed Section 27.1335 and incorporated its nese systems in separate rules that use diferent terminology nas resulted in som Page F - 54.1
AC 27.1337. § 27.1337 (Amendment 27-12) POWERPLANT INSTRUMENTS - (Paragraph (b) - FUEL QUANTITY INDICATOR).
a. Explanation. Section 27.1337(b) requires, in part, a means to indicate to the flight crew the quantity of usable fuel during flight in each tank. When two or more tanks are interconnected so that a failure of the system could cause fuel to become trapped in a fuel tank, the fuel quantity indicating system must provide the flight crew with the ability to determine the total effective amount of remaining usable fuel. Since the flight attitude of a rotorcraft may vary significantly with center of gravity (CG) and airspeed, several attitudes for calibration of the fuel gauge may be appropriate. Accordingly, the manufacturer should review the operational envelope (with respect to attitudes) and select at least three attitudes for gauge calibration with the provision that these attitudes should be useful to the intended operation of the rotorcraft and demonstrate gauge accuracy (within limits suggested herein). Selection of ground attitude for one of these conditions may, in some cases, be appropriate.
b. Procedures.
(1) Determine the rotorcraft pitch attitudes for most forward and most aft CG at a median gross weight and at an airspeed of 0.9 VNE or 0.9 VH, whichever is less. The mean attitude of the extremes defined above, further adjusted for lateral CG effects, if necessary, define the rotorcraft attitude for fuel gauge calibration.
(2) After establishing the calibration attitude, the requirements of § 27.1337(b) can be accomplished. The aircraft should be placed in the calibration attitude. Add fuel to the filler neck spillover level. Defuel the aircraft in increments corresponding to fuel gauge increment markings or at least 10 increments until gauge zero is obtained.
Precautions should be taken during this step to be sure that the fuel transmitter is sensing fuel level and not simply reflecting a physical "STOP" or end point in the system range. The fuel remaining in the tank below the "ZERO" mark must not be less than that amount determined by flight testing under § 27. 959. (Otherwise, the zero point must be adjusted upward.) The gauging system accuracy is acceptable when it meets a tolerance of +2 percent of the total useable fuel plus +4 percent of the remaining usable fuel at any gauge reading, provided that the gauge indicates zero fuel with unusable fuel in accordance with § 27.959 in the tank. (For a 100-gallon tank, this formula would allow a ‡6-gallon error at the full level, ‡4-gallon error at 50-gallon level, converging to a +2-gallon error at low fuel with the further provision that the zero mark accurately reflects unusable fuel.)
(3) Certain other aspects of a fuel gauging system need attention in order to minimize fuel exhaustion incidents:
(1) Gauge reading with the aircraft at ground attitude is frequently used by the crew in calculating range, weight and balance, and actual gross weight. Significant gauge errors in either direction during this reading can introduce hazards to the operation of the aircraft. If a calibration at this attitude indicates an unconservative error Page F - 54.2
AC 27-1B, Chg 4 3/31/2014 in excess of 6 percent of the gauge reading, corrective information should be applied adjacent to the fuel quantity gauge or be made available to the crew in other handbook data.
(ii) Flight during hover with maximum rearward wind may introduce significantly different fuel gauge readings. A check should be made to ensure that the gauge is either accurate or at least does not read high (unconservative) in this attitude.
(iii) If external loads are approved, the attitude for this type of operation should be considered.
(iv) Consistent with the requirements of § 27.1337 (b)(2), a separate fuel quantity indication is necessary for any interconnected fuel tank that has a flow control device, such as a fuel transfer pump or flapper valve, which could fail and trap fuel.
This requirement also applies to auxiliary fuel tanks. A sight gauge that is readable by the flight crew in flight may be acceptable for use with auxiliary fuel tanks.
(4) Fuel gauging system transmitters which are strictly volumetric measuring devices (float-actuated variable rheostats) introduce a gauge readout error of about 5 percent if calibrated with a fuel temperature of 0°C (32°F) and subjected to -55°C (-67°F) fuel or +55°C (131°F) fuel. This error may be minimized by calibrating the gauge with fuel temperature in the middle of the useful range (i.e., 15°C (59°F)). These types of gauging systems should be recalibrated during periodic major maintenance intervals. The manufacturer should be urged to include procedures and requirements for this in the "Instructions for Continued Airworthiness" for the model rotorcraft.
(5) Capacitance transmitters have become the standard for most modern fuel systems. These transmitters ordinarily need no temperature compensation since the fuel volume and the fuel dielectric constant vary inversely as temperature changes. The basic capacitance transmitter does not compensate for the different dielectric values to be expected with different type fuels. An add-on capacitance located where it will be submerged in fuel at all times can be devised to automatically compensate for other AC 27.1337A. § 27.1337 (Amendment 27-23) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 27-23 adds § 27.1337(e) that requires certain rotor drive system transmissions and gearboxes to be equipped with chip detector systems.
These detectors will sense and signal the presence of ferromagnetic particles to the flight crew. This amendment will improve the level of safety available with the installation of chip detector systems.
b. Procedures. All of the policy material pertaining to this section remains in effect. Additionally, the following information is added about chip detectors. The chip detectors should:
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(1) Indicate the presence of ferromagnetic particles in the transmission or gearbox;
(2) Be easily removable for inspection of the magnetic poles for metallic chips;
and,
(3) Prevent the loss of lubricant in the event of failure of the retention device for the removable portion of the chip detector (debris monitor).
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AC 27-1B, Chg 3 9/30/2008 [This page intentionally blank.} Page F - 54.5
SUBPART F - EQUIPMENT
ELECTRICAL SYSTEMS AND EQUIPMENT AC 27.1351. § 27.1351 (Amendment 27-13) ELECTRICAL SYSTEMS AND EQUIPMENT--GENERAL.
a. Explanation. With the advent of more sophisticated rotorcraft and operations under more critical conditions, such as IF and icing, it is essential that the electrical system be very carefully analyzed and evaluated to assure proper operation under any foreseeable operating condition and that hazards do not result from any malfunctions or failures.
b. Procedures.
(1) Electrical System Capacity. Rotorcraft electrical systems have grown in capacity, complexity, and impact on safety. This paragraph requires adequate electrical system capacity for safe operation of load circuits essential for safe operation at continuous rated power. If this capacity can be shown by electrical measurements, an electrical load analysis is not required. However, if the measurement method of compliance is chosen, data should be provided to show that the measured loads represent the worst case electrical loading for components or combination of components.
(1) Load circuits (systems) that are essential for safe operation are those systems necessary to maintain controlled flight and land safely and are generally those systems required to show compliance with the certification regulations. This includes most electrical utilization systems.
(ii) An electrical utilization system is a system of electrical equipment, devices, and connected wiring using electric energy to perform a specific aircraft function.
(ili) The specific utilization systems, which are necessary to maintain controlled flight and land safely, will vary with the type of rotorcraft and with the nature of operations. Examples of systems which may be essential are basic flight instruments, minimum navigation equipment, minimum radio communications, and flight control systems.
(2) Function.
Generating systems should be analyzed, inspected, or tested to ensure that no probable malfunction in the generating system or in the generator drive system may cause loss of service to systems essential for safe operation. A probable malfunction is any single electrical or mechanical component malfunction or failure that Page F - 55
AC 27-1B, Chg 3 9/30/2008 is likely to occur based on past service experience. This past service experience can include malfunction of components of previously approved rotorcraft, other aircraft, or qualitative analysis of similar components in rotorcraft applications. Analyses should be performed on the electrical power system emphasizing the exclusion of single point failures and the possibility of latent failures. Test methods should be developed that uncover latent faults. To identify latent failure causing events, consider the functions and the results of the safety analysis, as well as any failure detection monitor that may be incorporated. Evaluate the system using various combinations of unforeseen internal signal patterns and states. Also, evaluate the non-monitored functions by selecting test conditions that use every signal path and decision point between the input and output. Refer to MG-2 for electrical system test methods. These analyses should be extended to multiple malfunctions when:
(A) The first malfunction would not be detected during normal operation of the system, including periodic checks established at intervals which are consistent with the degree of hazard involved; or (B) The first malfunction would inevitably lead to other malfunctions.
(її) The generator drive system includes the prime movers (propulsion engines or other) and coupling devices such as gearboxes or constant speed drives.
(iii) Where crew corrective action is necessary:
(A) Adequate warning should be provided for any malfunction or failure requiring such corrective action;
(B) Controls should be located to permit such corrective action during any probable flight situation;
(C) If corrective action must be taken within a specified time to continue safe operation of the generating system, it should be demonstrated that such corrective action can be accomplished within the specified time during any probable flight situation; and (D) The procedure to be followed by the crew should be detailed in the Rotorcraft Flight Manual.
(iv) Chapter 11 of Advisory Circular 43.13-1A, "Acceptable Methods, Techniques, and Practices; Aircraft Inspection and Repair," includes guidance on installation of electrical systems (routing, separation, tying, clamping, j-box installations, etc.). Special emphasis should be placed on wire routing during the rotorcraft compliance inspection. Control wires to the rotorcraft's generators should be routed separately from generator output wiring. This should begin at the generator and continue to the voltage regulator. Additionally, wiring installation design should be Page F - 55.1
documented sufficiently to maintain configuration control for manufacturing and to assure that the electromagnetic characteristics remain the same as the certification sample.
(3) Generating System. When electrical power is needed for essential equipment, this paragraph requires at least one generator with adequate capacity for safe operation. Complete electrical failures have been caused by loss of voltage control in the voltage regulator. Overvoltage conditions can destroy electronic equipment. An acceptable method of overvoltage protection is the use of a separate overvoltage sensing relay to trip the generator off the line when overvoltage is detected. Another acceptable method is use of a voltage regulator with built-in overvoltage protection.
(4) Generator Ratings: Generator ratings are often the result of installation temperature limitations. The determination of these limitations, if any, is by testing the actual installation. The procedures for performing generator cooling tests are as follows:
Test Requirements.
(A) General. The applicant should contact the generator (alternator) manufacturer and obtain the maximum limits (continuous, 2 minute intermittent and 5 second transient 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.
(B) Instrumentation.
(1) Load Bank. A load bank will usually be necessary to load the test unit to the amperage limit for which approval is requested.
2) Ammeter. An ammeter should be provided with sufficient resolution to assure the amperage load is being maintained at the desired level.
(3) Temperature/Pressure Readouts. Readouts which are compatible with the temperature or pressure sensors installed in the test unit should be provided.
(4) Calibration Records. Calibration records should be available for all instrumentation.
(5) 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.
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AC 27-1B, Chg 3 9/30/2008
27. 1521(1), 27.10 41,27.103, 27.1045 (hug mement 27-19)., 27.1351, (D) 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). The tests are conducted to determine the maximum generator capacity that does not result in surpassing the limits given from the manufacturer. This is for a continuous rating, any credit for short time over current ratings must also be verified by the same methods, particularly for short time ratings longer than 5 seconds.
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(її) Test Procedures.
(A) Single-Engine Procedure.
(1) The cooling test is to be conducted during ground operation, climb-out, cruise, approach, and hover flight regimes.
2) All ground operational and in-ground-effect hover tests should be conducted in ambient winds of 5 knots or less.
(3) 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.
(4) 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.
(B) Multiengine Procedures. Conduct a generator cooling test in accordance with the following procedures:
1) All ground operational and in-ground-effect hover tests should be conducted in ambient winds of 5 knots or less.
(2) After engine start, load the instrumented generator to its proposed amperage limit and begin recording temperatures.
(3) A total of 30 minutes should be spent on the ground prior to takeoff.
This is to account for taxi and holding times.
(4) 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.
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