(13) Doppler Navigation.
(i) Doppler navigation system. Installed performance should be evaluated in accordance with AC 121-13.
(ii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(14) Audio Interphone Systems.
(i) Acceptable communications. These 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 (that is, emergency descent, and oxygen masks) with all engines running, all pulse equipment transmitting and all electrical equipment operating. If aural warning systems are installed, they should be evaluated, including distinguishing aural warnings when using headphones and with high air noise levels.
(ii) Electromagnetic Compatibility (EMC). With all systems operating during flight, verify, by observation, that no adverse effects are present in the required flight systems.
(15) Electronic Flight Instrument Systems. Refer to AC 23.1311-1.
(16) LORAN C Navigation Systems. Refer to AC 20-121A.
(17) Microwave Landing Systems. Reserved.
(18) Global Positioning System Navigation. Reserved.
2. § 23.1303 Flight and Navigation Instruments.
a. Free Air Temperature (FAT). Section 23.1303(a)(4) requires that reciprocating engine- powered airplanes of more than 6,000 pounds maximum weight and turbine-powered airplanes have a free air temperature indicator or an air temperature indicator that provides indications that are convertible to free air. The temperature pickup can be calibrated against a test pick-up of known characteristics, or by flying at various speeds at constant altitude, or by tower fly-by. This calibration is frequently done in conjunction with one or more of the airspeed calibration methods described in paragraph 23.1323 of this AC.
b Speed Warning Device. The production tolerances of the speed warning device required with § 23.1303(a)(5) must be set to minimize nuisance warnings. In considering this requirement, manufacturers should endeavor to reduce, lessen, or diminish such an occurrence to the least practical amount with current technology and materials. The least practical amount is that point at which the effort to further reduce a hazard significantly exceeds any benefit, in terms of safety, derived from that reduction. Additional efforts would not result in any significant improvements in reliability.
3. § 23.1305 Powerplant Instruments.
a. Explanation. Section 23.1305 is specific as to the powerplant instruments required for each type of installation. The requirement for specific instruments on specific airplanes should be determined by analysis of type design data prior to certification flight test.
b. Procedures. Verify proper functioning of each required instrument/indicator installed. If the creation of a required malfunction would require establishing a potentially hazardous condition in flight, proper functioning of these indicators may be verified by ground test.
c. Fuel Flowmeters. AC 23-17 covers the installation of fuel flowmeters in airplanes with continuous-flow fuel injection reciprocating engines.
4. § 23.1307 Miscellaneous Equipment. Reserved.
5. § 23.1309 Equipment, Systems, and Installations. This item is covered in AC
23.1309-1E.
Section 2. Instruments: Installation
1. § 23.1311 Electronic Display Instrument Systems. This item is covered in AC
23.1311-1.
2. § 23.1321 Arrangement and Visibility. Reserved.
3. § 23.1322 Warning, Caution, and Advisory Lights. Reserved.
4. § 23.1323 Airspeed Indicating System.
a. Explanation.
(1) Airspeed Indicator. An airspeed indicator is usually a pressure gauge that measures the difference between free stream total pressure and static pressure and is usually marked in knots. Pitot tubes for duplicate airspeed indicators are usually located on opposite sides of an aircraft fuselage but may be situated on the same side provided that they are separated by at least 14 inches (35.6 cm.) vertically.
(2) Air Data Computer Systems. Reserved.
(3) Definitions. Section 1.1 of part 1 of 14 CFR defines IAS, CAS, equivalent air speed (EAS), true air speed (TAS), and Mach number. These definitions include the terms position error, instrument error, and system error, which may need further explanation.
(i) Position Error. Position error is the total-pressure (pitot) and static-pressure errors of the pilot-static installation. By proper design, the total pressure error may be reduced to the point where it is insignificant for most flight conditions. NASA Reference Publication 1046 (refer to paragraph g gives various design considerations. The static pressure is more difficult to measure and the static position error may result in large airspeed and altitude errors.
(ii) Instrument Error. Instrument errors are errors inherent in the instrument for mechanical instruments. These errors are the result of manufacturing tolerances, hysteresis, temperature changes, friction, and inertia of moving parts. For electronic instruments, these errors are due to errors in the electronic element that converts pitot-static pressures into electronic signals. Instrument errors are determined for infl ight conditions in steady state conditions. Ground run system calib rations may require the consideration of internal instrument dynamics, electronic lags, and display filtering that would cause indication errors during takeoff acceleration.
(iii) System Error. System error is the combination of position error and instrument error.
(4) Temperatures. Static air temperature (SAT) a nd total air temperature (TAT) are not defined in section 1.1 of 14 CFR but may be significant in accurate calibration of airspeed systems. For stabilized values of pressure altitude and calibrated airspeed, TAS is a function of static air temperature. Reference (2) in section f of appendix 2 discusses the heating effect of the airflow on the temperature sensor and shows how to determine the recovery factor of the sensor.
Figure 7 of appendix 7 gives temperature ram rise, if the sensor recovery factor is known.
(5) System Calibration. The airspeed system is calibrated to determine compliance with the requirements of § 23.1323 and to establish an airspeed reference that is used in demonstrating compliance with other applicable regulations. The airspeed system may be calibrated using the speed course method, pacer airplane method, tower fly-by method, trailing cone method or either the trailing bomb or airspeed boom method or both. The method used will depend on the speed range of the airplane tested, the configuration, and the equipment available.
System calibration of the airspeed system is usually determined at altitudes below 10,000 feet.
For airplanes approved for flight above 31,000 feet, it is appropriate to verify validity of position errors at the higher operating altitudes. For airplanes where the static ports are located in close proximity to the propeller plane, it should be verified that sudden changes in power do not appreciably change the airspeed calibration. Additionally, for commuter category airplanes, § 23.1323(e) requires an airspeed calibration for use during the accelerate-takeoff ground run.
(6) Instrument Calibration. All instruments used during the test should be calibrated, and all calibration curves should be included in the type inspection report.
b. Speed Course Method. The speed course method consists of using a ground reference to determine variations between indicated airspeed and ground speed of the airplane.
Refer to appendix 9 for test procedures and a sample data reduction.
c. Trailing Bomb or Airspeed Boom Method. Refer to appendix 9 for procedures, test conditions, and a sample data reduction.
d. Pace Airplane Method. Refer to appendix 9 for test procedures.
e. Tower Fly-by. Refer to appendix 9 for test procedures.
f. Ground Run Airspeed System Calibration. The airspeed system is calibrated to show compliance with commuter category requirements of § 23.1323(c) during the acceleratetakeoff ground run and is used to determine IAS values for various V1 and VR speeds. Refer to appendix 9 for definitions, test procedures, and sample data reductions.
g. Other Methods. Other methods of airspeed calibration are described in NASA Reference Publication 1046, "Measurement of Aircraft Speed and Altitude," by W. Gracey, May 1980. Refer to appendix 9 for GPS method.
5. § 23.1325 Static Pressure System.
a. Definitions. Section 23.1323 defines several of the terms associated with the pitot static systems. Others may need further explanation.
(1) Altimeter. An altimeter is a pressure gauge that measures the difference between a sea level barometer pressure set on the instrument and static pressure, and indicates in units of feet.
(2) Static Error (error in pressure altitude). The error that results from the difference between the actual ambient pressure and the static pressure measured at the airplane static pressure source is called static error. Static error causes the altimeter to indicate an altitude that is different than actual altitude. It may also affect the errors in the airspeed indicating system.
b. Static System Calibration. The static system is calibrated to determine compliance with the requirements of § 23.1325. The static system may be calibrated by utilizing a trailing bomb, cone, or tower fly-by method. Alternately, for properly designed pitot systems, the pitot has minimal affect on the airspeed position error (dVc), as determined for § 23.1323. For these systems, static error (dh) may be calculated by the following equation:
2.5 VC 2 VC dh 0.08865dVC 1 0.2 ,ft. 661.5 where V c = calibrated airspeed, knots = ambient air density ratio dVc = airspeed position error
c. Test Methods. The methods specified for calibration of the airspeed indicating systems, including test conditions and procedures, apply equally for determining static error and error in indicated pressure altitude, and are usually determined from the same tests and data.
d. Tower Fly-by. The tower fly-by method is one of the methods that results in a direct determination of static error in indicated pressure altitude without the need for calculating from airspeed position error.
e. Procedures and Test Conditions for Tower Fly-by.
(1) Air Quality. Smooth, stable air is needed for determining the error in pressure altitude.
(2) Weight and C.G. Same as for calibrations of the airspeed indicating system.
(3) Speed Range. The calibration should range from 1.3 V SO to 1.8 VS1. Higher speeds up to VMO or VNE are usually investigated so that errors can be included in the AFM for a full range of airspeeds.
(4) Test Procedures.
(i) Stabilize the airplane in level flight at a height that is level with the cab of a tower, or along a runway while maintaining a constant height of 50 to 100 feet by use of a radio 150
altimeter. A ground observer should be stationed in the tower, or on the runway with an altimeter of known instrument error. Pressure altitude is recorded on the ground and in the airplane at the instant the airplane passes the ground observer.
(ii) Repeat step (i) at various airspeeds in increments sufficient to cover the required range and at each required flap setting.
(5) Data Acquisition. Data to be recorded at each test point:
(i) Airplane IAS;
(ii) Airplane indicated pressure altitude;
(iii) Ground observer indicated pressure altitude;
(iv) Radar altimeter indication (if flown along a runway);
(v) Wing flap position; and
(vi) Landing gear position.
(6) Data Reduction.
(i) Method.
(A) Correct indicated pressure altitude values for instrument error associated with each instrument.
(B) To obtain test pressure altitude, adjust the ground observed pressure altitude by the height read from the radar altimeter. No adjustment is required if the airplane was essentially the same level as the ground operator (tower cab). Static errors may be adjusted from test pressure altitude to sea level by the following:
dh (S.L.) = dh(TEST)(TEST) Where:
dh (TEST) = Difference in test pressure altitude and airplane pressure altitude with associated instrument errors removed.
(TEST) = ambient air density ratio.
(i) Plotting. Static error at sea level (dh (S.L.) ) should be plotted vs. test calibrated airspeeds.
(7) Required Accuracy. Section 23.1325(e) requires that the error in pressure altitude at sea level (with instrument error removed) must fall within a band of ±30 feet at 100 knots or less, with linear variation of ±30 feet per 100 knots at higher speeds. These limits apply for all flap settings and airspeeds from 1.3 VSO up to 1.8 VS1. For commuter category airplanes, the altimeter system calibration should be shown in the AFM.
6. § 23.1326 Pitot Heat Indication Systems. Reserved.
7. § 23.1327 Magnetic Direction Indicator. Reserved.
8. § 23.1329 Automatic Pilot System. This subject is covered in AC 23-17.
9. § 23.1331 Instruments Using a Power Supply. Reserved.
10. § 23.1335 Flight Director Systems. Reserved.
11. § 23.1337 Powerplant Instruments.
a. Explanation.
(1) Fuel Quantity Indicator. The indicator should be le gible and easily read without excessive head movement. The calibration units and the scale graduations should be readily apparent. Units should be consistent w ith AFM procedures and performance data.
(2) Auxiliary Tanks. A fuel quantity indi cator is not required for a small auxiliary tank that is used only to transfer fuel to another tank if the relative size of the tank, the rate of fuel flow, and operating instructions are adequate. The requirement for a separate quantity indicator should be determined by analysis of design data prior to flight test. The relative size of the tanks, intended use of the auxiliary tanks, complexity of the fuel system, and so forth, should be considered in determining the need for a fuel quantity indicator. If an indicator is not installed, flight manual procedures should ensure that once transfer of fuel is started, all fuel from the selected auxiliary tank can be transferred to the main tank without overflow or overpressure.
b. Procedures. Evaluate indicators for clarity and legibility. Review AFM for consistency of units and validity of procedures.
Section 3. Electrical Systems and Equipment
1. § 23.1351 General. Reserved.
2. § 23.1353 Storage Battery Design and Installation.
Reserved.
3. § 23.1357 Circuit Protective Devices. Reserved.
4. § 23.1361 Master Switch Switch Arrangement. This subpart requires a master switch arrangement to be installed. Confirm that the master switch arrangement is prominently located and marked. The master switch, in accordance with § 23.1355(e)(2), is considered to be an emergency control and should be colored red.
5. § 23.1367 Switches. Reserved.
Section 4. Lights
1. § 23.1381 Instrument Lights. Reserved.
2. § 23.1383 Landing Lights. Reserved.
Section 5. Safety Equipment
1. § 23.1411 General. Reserved.
2. § 23.1415 Ditching Equipment. Reserved.
3. § 23.1416 Pneumatic Deicer Boot System. Refer to AC 23-17.
4. § 23.1419 Ice Protection. Refer to AC 23-17 and AC 23.1419-2D.
Section 6. Miscellaneous Equipment
1. § 23.1431 Electronic Equipment. Section 23.1431(e) requires that the flight crewmembers will receive all aural warnings when any headset is being used. For those installations where not all warnings are provided through the radio/audio equipment, the manufacturers should demonstrate that all warnings will be heard and recognized when noise canceling headsets are used.
2. §23.1435 Hydraulic Systems. Reserved.
3. § 23.1441 Oxygen Equipment and Supply. Reserved.
4. § 23.1447 Equipment Standards for Oxygen Dispensing Units. Reserved.
5. § 23.1449 Means for Determining Use of Oxygen. Reserved.
6. § 23.1457 Cockpit Voice Recorders. Reserved.
7. § 23.1459 Flight Data Recorders. Reserved.
Chapter 6. Subpart G--Operating Limitations and Information Section 1. General 1 § 23.1501 General.
a. Explanation.
(1) Flight Crew Information. This section establishes the obligation to inform the flight crew of the airplane's limitations and other information necessary for the safe operation of the airplane. The information is presented in the form of placards, markings, and an approved AFM. Appendix 4 can be used to assist in determining which methods of presentation are required.
(2) Minimum Limitations. Sections 23.1505 through 23.1527 prescribe the minimum limitations to be determined. Additional limitations may be required.
(3) Information Presentation. Sections 23.1541 through 23.1589 prescribe how the information should be made available to the flight crew.
b. Procedures. None.
2. § 23.1505 Airspeed Limitations.
a. Explanation. This section establishes the operational speed limitations for safe margins below design speeds. For reciprocating engine-powered airplanes, there is an option.
They may either establish a never-exceed speed (VNE) and a maximum structural cruising speed (VNO) or they may be tested in accordance with § 23.335(b)(4), in which case the airplane is operated under a maximum operating speed concept (VMO/MMO). For turbine-powered airplanes, a VMO/MMO should be established. Tests associat ed with establishing these speeds are discussed under § 23.253, High speed characteristics.
b. Procedures. None.
3. § 23.1507 Maneuvering Speed. This regulation is self-explanatory.
4. § 23.1511 Flap Extended Speed. This regulation is self-explanatory.
5. § 23.1513 Minimum Control Speed. This regulation is self-explanatory.
6. § 23.1519 Weight and Center of Gravity. This regulation is self-explanatory.
7. § 23.1521 Powerplant Limitations. Reserved.
8. § 23.1523 Minimum Flight Crew. Reserved.
9. § 23.1523 Minimum Flight Crew.
a. Information. The following should be considered in determining minimum flight crew:
(1) Basic Workload Functions. The following basic workload functions should be considered:
(i) Flight path control;
(ii) Collision avoidance;
(iii) Navigation;
(iv) Communications;
(v) Operation and monitoring of aircraft controls;
(vi) Command decisions; and
(vii) Accessibility and ease of operation of necessary controls.
(2) Workload Factors. The following workload factors are considered significant when analyzing and demonstrating workload for minimum flight crew determination:
(i) The impact of basic airplane flight characteristics on stability and ease of flight path control. Some factors such as trimmability, coupling, response to turbulence, damping characteristics, control breakout forces and control force gradients should be considered in assessing suitability of flight path control. The essential elements are the physical effort, mental effort and time required to track and analyze flight path control features, and the interaction with other workload functions.
(ii) The accessibility, ease, and simplicity of operation of all necessary flight, power, and equipment controls, including emergency fuel shutoff valves, electrical controls, electronic controls, pressurization system controls, and engine controls.
(iii) The accessibility and conspicuity of all necessary instruments and failure warning devices such as fire warning, electrical system malfunction, and other failure or caution indicators. The extent to which such instruments or devices direct the proper corrective action is also considered.
(iv) The complexity and difficulty of operation of the fuel system, with particular consideration given to the required fuel management schedule necessitated by c.g., structural, or other airworthiness considerations. Additionally, the ability of each engine to operate 156
continuously from a single tank or source that is automatically replenished from other tanks if the total fuel supply is stored in more than one tank.
(v) The degree and duration of concentrated mental and physical effort involved in normal operation and in diagnosing and coping with malfunctions and emergencies, including accomplishment of checklist, and location and accessibility of switches and valves.
(vi) The extent of required monitoring of the fuel, hydraulic, pressurization, electrical, electronic, de-icing, and other systems while en route. Also, recording of engine readings, and so forth.
(vii) The degree of automation provided in the event of a failure or malfunction in any of the aircraft systems. Such automation should ensure continuous operation of the system by providing automatic crossover or isolation of difficulties and minimize the need for flight crew action.
(viii) The communications and navigation workload.
(ix) The possibility of increased workload associated with any emergency that may lead to other emergencies.
(x) Passenger problems.
(3) Kinds of Operation Authorized. During minimum crew determination, consideration should be given to the kinds of operation authorized under § 23.1525. Inoperative equipment could result in added workload that would affect minimum crew. It may be determined that, due to minimum crew workload considerations, certain equipment must be operative for a specific kind of operation.
b. Acceptable Techniques.
(1) General.
(i) A systematic evaluation and test plan should be developed for any new or modified airplane. The methods for showing compliance should emphasize the use of acceptable analytical and flight test techniques. The crew complement should be studied through a logical process of estimating, measuring, and then demonstrating the workload imposed by a particular flight deck design.
(ii) The analytical measurements should be conducted by the manufacturer early in the airplane design process. The analytical process that a given manufacturer uses for determining crew workload may vary depending on flight deck configuration, availability of a suitable reference, original design or modification, and so forth.
(2) Analytical Approach.
(i) A basis for deciding that a new design is acceptable is a comparison of a new design with a previous design proven in operational service. By making specific evaluations and comparing new designs to a known baseline, it is possible to proceed with confidence that the changes incorporated in the new designs accomplish the intended result. When the new flight deck is considered, certain components may be proposed as replacements for conventional items, and some degree of rearrangement may be contemplated. New avionics systems may need to be fitted into existing panels, and newly automated systems may replace current indicators and controls. As a result of these evolutionary charac teristics of the flight deck design process, there is frequently a reference flight deck design, which is usually a conventional airplane that has been through the test of operational usage. If the new design represents an evolution, improvement attempt, or other deviation from this reference flight deck, the potential exists to make direct comparisons. While the available workload measurement techniques do not provide the capacity to place precise numbers on all the relevant design features in reference to error or accident potential, these techniques do provide a means for comparing the new proposal to a known quantity. Service experience should be res earched to assure that any existing problems are understood and not perpetuated.
(ii) After studying a new component or arrangement and exercising it in practical flight scenarios, a test pilot may not be able to grade that design in finer workload units than "better" or "worse than." If the pilot can say with reliability and confidence that it is or is not easier to see a display or to use an augmented control system than to use a functionally similar unit of a reference design, then these "better" or "worse than" judgments provide substantial evidence that workload is or is not reduced by the innovation. These “better” or “worse than” judgments should be corroborated b a reasonable sample of qualified pilots over various assumed flight regimes.
(A) If an early subjective analysis by FAA flight test personnel shows that workload levels may be substantially increased, a more in-depth evaluation of flight testing may be required to prove acceptability of the increased workload. In this case, there should be available workload latitude in the basic flight deck design to accommodate the increase.
(B) If the new design represents a "revolutionary" change in level of automation or pilot duties, analytic comparison to a reference design may have lessened value.
Without a firm database on the time required to accomplish both normally required and contingency duties, more complete and realistic simulation and flight testing will be required.
(3) Testing.
(i) In the case of the minimum crew determination, the final decision is reserved until a panel of experienced pilots has flown the airplane, trained and qualified in the airplane.
The training should be essentially that required for a type rating. When the applicant seeks single pilot approval, the evaluation pilots should be experienced and proficient in single pilot operations. Section 23.1523 contai ns the criteria for determining the minimum flight crew.
These criteria contain basic workload functions and workload factors.