(1) The following instruments are considered to be flight instruments:
(i) Airspeed indicator.
(ii) Sensitive altimeter.
(iii) Free-air temperature indicator.
(iv) Nontumbling gyroscopic bank and pitch indicator.
(V) Gyroscopic rate-of-turn indicator with an integral slip-skid indicator.
(vi) Rate-of-climb (vertical speed) indicator.
(2) The remaining instruments are navigation instruments.
(3) If a speed warning device is required to be included as part of the rotorcraft design, it must meet the performance requirements given in the rule. In addition, the evaluation of the acceptability of the aural warning (i.e., this warning differs distinctively from aural warnings used for other purposes) should be accomplished by flight test personnel as part of their overall cockpit evaluation.
(4) Electronic Flight Instrument Systems (EFS).
(i) Explanation. The increased use of microprocessor technology in avionic systems has resulted in the use of computer-generated graphics to replace conventional electromechanical instruments which are used for the display of flight information required by § 29.1303(1), (g), and (h). For IFR certified aircraft, the EFIS usually is used for the display of the magnetic gyro-stabilized direction indicator (slaved compass system.) These computer-generated graphics are usually displayed on small multicolor-shadow-mask cathode ray tubes (CRT) and replace the horizontal situation indicator (HSI) and the attitude direction indicator (ADI). This discussion presumes that the EFIS for which approval is sought meets the general requirements of an EFIS for a transport category airplane with regard to color, symbology, operation, and so forth.
This paragraph along with some others in this document principally highlights the areas which are peculiar to the installation in a transport category rotorcraft. A discussion of the flight director function performed by the EFIS is given in Paragraph 640. A discussion of the location of the displays is contained in Paragraph 632. A discussion of the requirements for an EFIS in a transport category airplane is contained in AC 25-11, Transport Category Airplane Electronic Display Systems, dated September 16, 1987.
(i¡) Procedures.
940 Par 618
(A) System Components. The system components require qualification testing to determine that their design is acceptable, free from hazards, and suitable to their airborne environment. Generally the components of the EFIS should meet the requirements of TSO C-113.
(1) Environmental Qualification. The EFIS hardware must be shown to be suitable to its airborne environment. A desirable way to qualify the system component is to obtain approval to the appropriate TSO. If the equipment is not TSO approved, it should be shown via testing that it complies with the requirements of SAE Document AS-8034. This will include testing in accordance with the appropriate categories of the latest revision of RTCA Document DO-160, JEDEC Publication No. 64D (Protection from lonizing Radiation), and UL Document No. 1418 (Impact Implosion Test).
(2) Software. The embedded software should be qualified to an appropriate standard. The software level is contingent on the worst case criticality of the function it performs. As an example, the display of incorrect roll and/or pitch by an EFIS ADI instrument is a hazardous or "critical" malfunction. The software should be designed to provide adequate consideration for this factor (reference Paragraph 621).
A similar consideration is required for altitude and airspeed.
(B) System Installation Consideration.
(1) Display Chromaticity and Luminance. The chromaticity and luminance of the displays should be determined to be acceptable for all critical cockpit lighting conditions which are expected in service. An expanded discussion of these characteristics may be found in AC 25-11.
(2) Temperature Survey to Determine Proper Cooling of EFIS Components.
(i) Equipment Requiring Cooling Test. As with any avionic equipment, good engineering judgment may deem that all components of the EFIS should have an in-flight temperature survey performed to ascertain that the thermal environmental tolerance of the system components is not exceeded. Usually, the following general guidelines may be used to aid in determining when an in-flight temperature survey is warranted.
(a) Components which contain a CRT require the temperature survey outlined in 618b(4)(ii)(B)(1) (i).
(b) Equipment which does not contain a CRT, but is specified by the manufacturer to require forced air cooling (by an airframe mounted system), usually requires a temperature survey.
Par 618 941
(c) Equipment which does not contain a CRT and is not specified as requiring forced air cooling may usually have its critical thermal environment substantiated by laboratory testing.
(…) Temperature Survey Testing. The temperature tests for the EFIS units should consist of a short-term test of approximately 30 minutes which accounts for an aircraft which has heat-soaked on the ramp. A factor of 25° F should be added to the maximum corrected temperature to account for "greenhouse effect." A long-term test should be accomplished at various altitudes and limiting (low and high) airspeeds. All avionic equipment should be turned on during this test, and the cockpit panel lights should be operated at full intensity. The environmental control unit (ECU) or air-conditioning system should not be operating during these tests; however, any windows or vents which are part of the "basic" TC rotorcraft may be utilized to ventilate the pilot's stations. Both these tests should be corrected to the maximum temperature for which the rotorcraft is certified and a standard lapse rate for altitude as specified in this AC. If an airframe cooling system is necessary to keep the display units within acceptable temperature limits, then the pilots) must be made aware of a failure or malfunction of this cooling system. Some type of cockpit visual annunciation with the capability to perform a preflight test is usually utilized to fulfill this requirement.
(3) System Reliability.
Failure of the EFIS to perform a required function which results in the reversion to standby instruments or requires the use of abnormal procedures should be shown to be improbable (for Category A rotorcraft, reference § 29.1309(b)).
(i1) For IFR operations, Appendix B of Part 29, paragraph VIlI(b)(5)(ii), requires that the equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the instruments will remain available to a pilot, without additional crewmember action, after any single failure or combination of failures that is not shown to be extremely improbable. The display of attitude, altitude, or airspeed is individually "essential to the safety of flight," and, therefore, the loss of all attitude display, all airspeed information, or all altitude information to the pilots) should be extremely improbable. Also, any malfunction or failure of the EFIS which would result in the display of simultaneous incorrect display of this critical information should be extremely improbable. In view of the relatively new technology embodied in the EFIS, the conventional technology electromechanical standby attitude indicator, with its independent power supply, should be retained.
(4) Symbology and Function. When assessing the acceptability of the EFIS, consideration should be given to the effect of the loss of one of the CRT color guns. This type of failure is especially a factor in determining the acceptability of the installation for single-pilot operation.
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(5) Circuit Protective Devices. All circuit protective devices for the EFIS and related, required interface units should be placed in the cockpit where they can be reset by one of the pilots without leaving the pilot seat.
c. Standby Instruments. The EFIS which have been approved on transport category rotorcraft at this time have only presented the critical function of attitude display. A specific requirement for a standby attitude instrument is contained in Appendix b or Part 29. his requirement is usually satistied by an electromechanical panel-mounted gyro with an independent power source. This type of installation is what was envisioned by the authors of Appendix B. Because of the mature technology of this type of standby attitude indicator, certain aspects of the EFIS installation have not been an area for concern. If, however, a total commitment of critical display functions is made to the "glass" technology, such that the standby attitude instrument requirement is satisfied by a software based CRT system, then several major areas of concern will be raised. Among these are the electromagnetic vulnerability of the system (protection from the effects of lightning and high energy radio frequency fields) and software. The certifications of EFIS with an electromechanical standby attitude indicator have not considered loss of function critical from a software aspect. (reference Paragraph 621 for a discussion of software qualification.)
619. § 29.1305 (Amendment 29-10) POWERPLANT INSTRUMENTS.
a. Explanation. This section specifies those instruments which are required for reciprocating and turbine engine installations. It also provides instrumentation requirements for operating rotorcraft in Category A or Category B. These instruments will provide the pilot with essential data to determine operational status of critical components and select desired performance conditions.
b. FAR 29.1305(a)(4), (a)(6), and (a)(9) requires a warning device for low tuel, gearbox oil pressure, and transmission oil temperature. An indicator/gage is not acceptable for use as a warning device since the indicator/gage is not a primary instrument and therefore is not actively monitored.
c. There are advanced display systems that take advantage of microprocessor power by integrating the processing of several parameters. These systems have to date been referred to as Engine Caution Advisory Systems (ECAS) or as Integrated Instrument Display Systems (IIDS) and possibly other variations of these names.
These systems typically integrate propulsion instruments, fuel quantity indication, and caution and warning system into a single display system. In traditional designs the powerplant instruments, fuel quantity display, and the caution and warning system are independent from each other. The integration of these systems/indicators eliminates their independence from one another and increases the probability of loss of more than one indicator/system as a result of a single fault or malfunction. Redundant design is generally applied to compensate for the loss of independence.
Par 618 943
(1) This integration and resultant mitigation of independence can result in an increased opportunity for common mode failures. Approval of the compensating features is elevated in importance as it is this aspect that allows the concept to be acceptable and subsequently certifiable. The loss of all displayed information or erroneous information should be considered for determination of worst case criticality.
With this determination of criticality, the design can be evaluated to see that it meets the minimum associated level of design assurance. Additionally, due to space limitations, some systems employ "page over" features that may have some difficulty displaying the required information when needed and human factors aspects must be considered.
(2) The instrument display system must be investigated and found to be acceptable under both normal and emergency conditions, must perform its intended function under foreseeable operating conditions, and must be designed to minimize the hazards in the event of probable malfunction or failure.
(3) It must be shown that there is appropriate redundancy to provide adequate compensation for the loss of independence in the system. If a multi-page system is employed, it must be shown that needed information is displayed when required.
Specific issues that must be addressed to assure compliance with the minimum safety standards are as follows:
The level of most severe hazard must be determined.
Equivalent reliability and software design assurance to the determined criticality level must be shown.
(ili) Pretest capability must be provided for the warning and caution system to preclude an associated latent failure.
(iv) Human Factors (reference §§ 29.1321 and 29.1322).
d. Additional rules to be consulted when determining the powerplant instrument installation design are §§ 29.1321, Arrangement and visibility; 29.1337, Powerplant instruments; 29.1381, Instrument lights; 29.1543, Instrument markings: General;
29.1549, Powerplant instruments; 29.1551, Oil quantity indicator; and 29.1553, Fuel quantity indicator.
619A. § 29.1305 (Amendment 29-26) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-26 revises, edits, and adds new powerplant instrument requirements. Section 29.1305(a)(4) was revised to require a low fuel warning device for each tank that can be used to feed an engine. The amendment allows a longer time between warning actuation and fuel exhaustion and requires the low fuel warning device to be independent of the normal fuel quantity indication system.
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7/30/97/ AC 29-2B Section 29.1305(a)(17) was changed to extend its application to all rotorcraft (not just those with turbine engines) and to require an indication to the crew of the degree of filter blockage as it relates to the fuel flow requirements in § 29.955.
Section 29.1305(a)(19) was revised to require function indicators only for fuel heaters that can be selected or are controllable. A new paragraph (a)(20) was added to § 29.1305 that combined identical requirements for fuel pressure indicators in Paragraphs (b)(2) and (c)(2) and modified the applicability of these requirements to only those fuel systems with devices or components that could adversely affect fuel pressure at the engine, if they fail. It also eliminates the requirement for fuel pressure indicators in fuel systems, such as suction or gravity feed systems, which do not incorporate pumps or filters. A new § 29.1305(a)(21) was added that requires a warning device to indicate to the flight crew the failure of any fuel pump that is required to supply adequate fuel flow to the engine according to § 29.955; such indication is not required for fuel pumps which are demonstrated to be only necessary for engine starting.
Section 29.1305(a)(22) adds a requirement for a warning or caution device to alert the flight crew when particles are detected by the chip detector required by § 29.1337 (d). A new § 29.1305(a)(23) added a requirement for powerplant instruments or warning devices for auxiliary power units installed in rotorcraft.
b. Procedures. The requirement and purpose for each instrument is self-explanatory in the amendment. Other sections that should be considered when designing powerplant instruments are listed in Paragraph 619.
619B. 29.1305 (Amendment 29-34) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-34 added §§ 29.1305(a)(24) and 29.1305(a)(25) to provide for 30-second/2-minute OEl power ratings.
(1) Section 29.1305(a)(24) adds the requirement that a device or means be provided to alert the crew of the use of the 30-second and 2-minute OE power level.
The crew should be alerted when the 30-second or 2-minute interval begins and when the time interval ends. The amount of time spent at the 30-second or 2-minute OEl power levels is at the crew's discretion, unlike the other limits for 30-second OLl that are set by an automatic control required by Paragraph 29.1143. The purpose for providing the time interval alerts and automatically controlling the 30-second OEI limits is to free the crew from monitoring the engine instruments during critical phases of flight caused by the loss of an engine.
(2) Section 29.1305(a)(25) adds the requirement for a device to record the usage and the amount of time spent at the 30-second OEI power level and the amount of time spent at the 2-minute OEI power level. The information recorded by this device is for the use of the ground crew to determine if maintenance actions/inspections are to be conducted.
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b. Procedures. For the purpose of complying with FAR 29.1305(a)(24) and
29.1305(a)(25), the 2-minute OE power level is considered to be achieved whenever one or more of the operating limitations applicable to the next lower OE power rating are exceeded. The 30-second OEl power level is considered to be achieved whenever one or more of the operating limitations applicable to the 2-minute OEl power rating are exceeded.
(1) A review of the method to meet the requirements of § 29.1305(a)(24) should be conducted by flight test personnel. A determination should be made as to whether the method used to alert the crew of 30-second or 2-minute OE power usage can be recognized and understood by the crew.
(2) To meet the requirements of Section 29.1305(a)(25), a device should be installed on the engine or the airframe to record the time and each usage of 30-second and 2-minute OEI power levels. The information on the time and usage of 30-second and 2-minute OEI power should be recoverable from the recording device by ground personnel. The device should not be capable of being reset in flight and should only be capable of being reset by ground personnel. Prior to each flight this device should be capable of being checked for proper operation and to determine if 30-second or 2-minute OEI power levels were used during the previous flight.
619C. § 29.1305 (Amendment 29-40) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-40 added section 29.1305(a)(6) to require an oil pressure indicator for each pressure-lubricated gearbox. Paragraphs (a)(6) through
(a)(25), prior to this amendment, have been redesignated as Paragraphs (a)(7) through
(a) (26).
b. Procedures. In addition to providing an oil pressure indicator for each pressure-lubricated gearbox, the guidance material of the previous 619 paragraphs continues to apply.
620. § 29.1307 (Amendment 29-12) MISCELLANEQUS EQUIPMENT.
a. Explanation. This rule provides a listing of several items of required miscellaneous equipment. Each item seems to be self-explanatory except the one requiring a master switch arrangement for electrical circuits other than ignition. The purpose of a master switch arrangement is to allow rapid removal of all bus loads from sources of electrical power in an emergency situation.
b. Procedures. When reviewing possible solutions to the master switch arrangement requirement, the following considerations should be included.
(1) System separation. Since wiring from each electrical system will be brought in close proximity to each other, extra care should be taken to maintain some 946 Par 619B
separation. As examples, common connectors, common grounds, and common wire routing should be avoided.
(2) Installation of switches. The single switch should be avoided since it introduces the possibility of a single failure turning off the entire electrical system. One solution that is commonly used provides a close grouping of the switches such that the pilot can easily reach all switches and turn them all off with one action. This solution requires a cockpit evaluation to ensure the installation will be suitable for different hand sizes. Another solution involves a gang bar that can be moved with a single motion to turn off all sources. This solution has been found to be acceptable in several instances.
Other solutions should be evaluated on their own merits, and the primary emphasis should be on maintaining some minimum system separation and conducting a cockpit evaluation by flight test personnel.
621. § 29.1309 (Amendment 29-40) EQUIPMENT, SYSTEMS. AND INSTALLATIONS.
a. Types of Equipment. This regulation covers, but is not limited to electrical, pneumatic, and hydraulic power sources, associated distribution, and corresponding utilization systems.
b. Environmental Qualification.
(1) Laboratory Tests.
Environmental Standards. In order to assure that the components/systems under consideration will function properly when exposed to adverse environments, they should be tested in the laboratory under a simulated adverse environment. If a TSO exists and it is appropriate in environmental range and performance for an equipment installation, it is preferable the equipment be TSO approved. If there is no applicable TSO or an existing TSO does not provide for a sufficiently adverse environment, the latest revision of the Radio Technical Commission for Aeronautics (RTCA) document DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.
(ii) Adverse environmental variables for all types of required and critical equipment include, but are not necessarily limited to temperature, humidity, vibration, shock, altitude, overpressure, and power source transients.
(iii) For electrical/electronic equipment, adverse environmental variables include all of (b) above plus overvoltage and undervoltage. Electronic equipment should also be tested for electromagnetic interference (EMI). These tests should include both emission and susceptibility evaluations of both conducted and radiated EMI.
Par 620 947
(iv) Explosion Tests. Those items of electrical/electronic equipment that are to be located in areas subject to flammable fluids and vapors, as a result of any single probable maltunction or tailure, including failure of couplings or lines should be tested as an ignition source. These tests consist of normal operation of the equipment in a physically contained explosive atmosphere. The explosion test procedure in the latest revision of DO-160 will satisfy this requirement. Paragraph 362 of this AC provides further guidance on safety from explosion. If another standard is used that is at least as good as the latest revision of DO-160, it may also be accepted to satisfy this requirement.
(2) Installed Environmental Tests. After the environmental ratings of the components/systems have been established, it should be assured that as installed, these ratings will not be exceeded. Normally, installed equipment need not be instrumented and tested in flight nor is it necessary to instrument the compartment or rack where the equipment is installed. Satisfactory environment and equipment compatibility are assured by selection of the proper environmental category of laboratory tests. The category is determined by the type of aircraft (reciprocating or turbine) and flight envelope (altitude and temperature). Exceptions to normal installations are (a) Alternator/generator cooling, where radiated and conducted heat is almost always uncertain, also cooling air temperatures and flow rates are uncertain;
(b) Where flight tests reveal excessive instrument panel vibration. In this case, the panel should be instrumented, tested, and, if necessary, design improvements made;
and (c) Any other cases where good engineering judgment and application of sound engineering principles indicate a high likelihood that the installed environment is more severe than the equipment is capable of operating within.
Temperature Tests.
(A) Temperature tests may be accomplished by instrumenting the installed equipment environment with a recorder that provides a permanent record of time, altitude, and temperature. The pertinent temperature should be recorded as the rotorcraft is operated throughout its altitude range, including ground operation. The maximum and minimum temperatures recorded should be corrected degree for degree to assure the equipment under test remains within its temperature rating while the rotorcraft operates throughout its approved ambient temperature envelope. (For generator/alternator cooling test procedures, refer to Paragraph 778 of this AC.)
Section 29.1043, Paragraph (b) requires the maximum approved operating OAT to be at least 100° F for powerplant-mounted accessories such as starter generators, vacuum pumps, etc. Due to the impracticality of the 100° F hot day temperature limit, rotorcraft systems mounted on the powerplant are normally evaluated for at least 115° F hot day sea level conditions with corresponding 3.6° F/1,000-foot correction. The maximum hot day OAT at sea level must be specified in the rotorcraft flight manual. Section 29.1043, Paragraph (b) is the regulatory basis for the lapse rate of 3.6° F/1,000 feet. This lapse rate should be applied regardless of the hot day sea level temperature the applicant chooses to certify for operation.
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(B) The § 29.1043 Paragraph (b) maximum ambient temperature definition should not be confused with operating temperatures in closed areas. Closed equipment rack areas can easily reach temperatures of 140° F when sitting on the ramp in the southern United States in midsummer. Normally, proper selection of the altitude temperature category in the latest revision of DO-160 will assure compliance.
(C) In some cases, the equipment manufacturer furnishes temperature limits for internal critical parts. For example, brushes, bearings, or field windings on DC generators. In these cases it is better to record the critical component temperature rather than equipment or equipment environment temperature.
(D) The following will illustrate an acceptable high temperature evaluation method:
TOAT MAX = Maximum outside air temperature at which temperature tests are conducted.
MAX = Maximum temperature to which the installed equipment has been tested in the laboratory.
TTEST MAX = Maximum installed equipment temperature recorded during tests.
ToR = The high reference outside air temperature. It varies with altitude starting at the highest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000 foot altitude. It can be no less than 100° F (reference § 29.1043(b));
however, it can be as high as the applicant wants.
TH MAR = Temperature margin between the maximum equipment temperature substantiated in the laboratory and the maximum installed equipment temperature when the rotorcraft is operating in the highest available OAT and approximately corrected at the altitude under consideration. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.
TH MAR = TMAX - (TTEST MAX + (TORH - TOAT MAX)) Example #1: Assume the applicant is seeking approval for rotorcraft operation at the lowest acceptable OAT, at sea level, of 100° F and TMAX for Generator Brush = 295° F at maximum load current throughout the altitude range. In-flight test data are:
Par 621 949
Measured Altitude (ft. MSL) Brush Temp (ITEST MAX) OAT = TOAT MAX sea level 275° F 90° F 5,000 270° F 80° F 10,000 285° F 60° F 15,000 294° F 42° F 20,000 290° F 20° F First, ToR must be calculated for each altitude test point.
@ sea level, TORH = 100° F @ 5,000 ft., ToR = 100° F -5,000 ft., × 3.6°/1,000 ft. = 82° F @ 10,000 ft., ToRn = 100° F - 10,000 ft. x 3.6/1,000 ft. = 64° F @ 15,000 ft., ToR = 100° F - 15,000 ft. x 3.6/1,000 ft. = 46° F @ 20,000 ft., ToR = 100° F - 20,000 ft. × 3.6°/1,000 ft. = 28° F Then at sea level:
TH MAR = 295 - (275 + (100 - 90)) At 5,000 feet:
TH MAR = 295 - (270 + (82 - 80)) = 23° F At 10,000 feet:
TH MAR = 295 - (285 +(64-60)) = 6° F At 15,000 feet:
TH MAR = 295 - (294 + (46 - 42)) = -3° F At 20,000 feet:
TH MAR = 295 - (290 + (28-20)) = -3° F Since TH MAR comes out negative at the 15,000- and 20,000-foot points, the generator fails. It will be necessary for the applicant to reduce the maximum load current, improve cooling, or otherwise change the design to assure the generator is operating within its approved temperature limit of 295° F.
(E) In most cases, the equipment is laboratory tested to minimum temperatures as severe as that of the rotorcraft's maximum certified altitude on a minimum temperature day. Therefore, unless equipment minimum temperature is 950 Par 621
affected by refrigeration or other temperature reducing environments, actual installed instrumented minimum temperature tests are unnecessary. If low temperature evaluation is necessary for the installed equipment, the following is an acceptable method:
ToAT MIN = Minimum outside air temperature at which temperature tests are conducted.
TMIN = Minimum temperature to which the installed equipment has been tested in the laboratory.
TTEST MIN = Minimum installed equipment temperature recorded during tests.
TORL = The low reference outside air temperature. It varies with altitude starting at the lowest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000-foot altitude.
T1 MAR = Temperature margin between the minimum equipment temperature substantiated in the laboratory and the minimum installed equipment temperature. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.
T1 MAR = -(TMIN - (TTEST MIN + (TORL - TOAT MIN))) NOTE: This equation assumes all temperatures are negative. It is necessary to place a (-) in front of the right side of the equation in order to convert the T1 MAR Value to the conventional positive answer for acceptance and a negative answer for rejection. Temperature in the 0 to 32° F range can be handled by conversion to the centigrade scale.
Example #2: Assume the applicant is seeking a low temperature operating limit at sea level of -25° F. Assume the hydraulic control cylinder has been substantiated in the laboratory to operate at a cylinder temperature of -40° F.
The in-flight test data are:
Measured Altitude (ft. MSL) Cylinder Temp (ITEST MIN) OAT = TOAT MAX sea level 0° F -25° F 5,000 -9° F -45° F 10,000 -21° F -59° F 15,000 -32° F -65° F 20,000 -40° F - 69° F Par 621 951