(vii) Orbiting. At an altitude of 2,000 feet above the terrain, at holding pattern speeds appropriate for the type of airplane and with the landing gear extended, fly at least 15 degree sectors of left and right 35 n.m. orbital patterns around the DME facility. The DME should continue to track with no more than one unlock, not to exceed one search cycle, in any 5 miles of orbited flight.
(viii) Approach. Make a normal approach at an actual or simulated field with a DME.
The DME should track without an unlock (station passage expected).
(ix) DME Hold. With the DME tracking, activate the DME hold function. Change the channel selector to a localizer frequency. The DME should continue to track on the original station.
(9) Transponder Equipment.
(i) Signal Strength. The ATC transponder system should furnish a strong and stable return signal to the interrogating radar facility when the airplane is flown in straight and level flight throughout the airspace within 160 n.m. of the radar station from radio line of sight to within 90 percent of the maximum altitude for which the airplane is certificated or to the maximum operating altitude. Airplanes to be operated at altitudes not exceeding 18,000 feet should meet the above requirements to only 80 n.m.
(ii) Single Site Tracking. Special arrangements should be made for single-site tracking.
ATC coverage includes remote stations and, unless single-site is utilized, the data may be invalid.
(iii) Dropout Times. When the airplane is flown within the airspace described above, the dropout time should not exceed 20 seconds in the following maneuvers:
(A) In turns at bank angles up to 10 degrees;
(B) Climbing and descending at normal maximum climb and descent attitude; and
(C) Orbiting a radar facility.
(iv) Climb and Distance Coverage.
(A) Beginning at a distance of at least 10 n.m. from and at an altitude of 2,000 feet above that of the radar facility and using a transponder code assigned by the ARTCC, fly on a heading that will pass the airplane over the facility. Operate the airplane at its normal maximum climb attitude up to within 90 percent of the maximum altitude for which the airplane is certificated, maintaining the airplane at a heading within 5 degrees from the radar facility. After reaching the maximum altitude for which the airplane is certificated, fly level at the maximum altitude to 160 (or 80) n.m. from the radar facility.
(B) Communicate with the ground radar personnel for evidence of transponder dropout. During the flight, check the "ident" mode of the ATC transponder to ensure that it is performing its intended function. Determine that the transponder system does not interfere with other systems aboard the airplane and that other equipment does not interfere with the operation of the transponder system. There should be no dropouts for two or more sweeps.
(v) Long-Range Reception. Perform two 360 degree turns, one to the right and one to the left, at bank angles of at least 10 degrees with the flight pattern at least 160 (or 80)
n.m. from the radar facility. During these turns, the radar display should be monitored and there should be no signal dropouts (two or more sweeps).
(vi) High-Angle Reception. Repeat the flight pattern and observations of (d) above at a distance of 50 to 70 n.m. from the radar facility and at an altitude of at least 90 percent of the maximum operating altitude. There should be no dropout (two or more sweeps).
Switch the transponder to a code not selected by the ground controller. The airplane secondary return should disappear from the scope. The controller should then change his control box to a common system and a single slash should appear on the scope at the airplane's position.
(vii) High-Altitude Cruise. Fly the airplane within 90 percent of its maximum certificated altitude or its maximum operating altitude beginning at a point 160 (or 80)
n.m. from the radar facility on a course that will pass over the radar facility. There should be no transponder dropout (two or more sweeps) or "ring-around."
(viii) Holding and Orbiting Patterns.
(A) At an altitude of 2,000 feet or minimum obstruction clearance altitude (whichever is greater) above the radar antenna and at holding pattern speeds, flaps and gear extended, fly one each standard rate 360 degree turn right and left at a distance of approximately 10 n.m. from the ARSR facility. There should be no signal dropout (two or more sweeps).
(B) At an altitude of 2,000 feet or minimum obstruction clearance altitude (whichever is greater) above the radar antenna and at holding pattern speeds appropriate for the type of airplane, fly 45 degree sectors of left and right 10 n.m.
orbital patterns around a radar facility with gear and flaps extended. There should be no signal dropout (two or more sweeps).
(i) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(10) Weather Radar.
(i) Bearing Accuracy. The indicated bearing of objects shown on the display should be within ±10 degrees of their actual relative bearing. Verify that as airplane turns to right or left of target, the indicated display moves in the opposite direction. Fly under conditions that allow visual identification of a target, such as an island, a river, or a lake, at a range of approximately 80 percent of the maximum range of the radar. When flying toward the target, select a course that will pass over a reference point from which the bearing to the target is known. When flying a course from the reference point to the target, determine the error in displayed bearing to the target on all range settings. Change heading in increments of 10 degrees and determine the error in the displayed bearing to the target.
(ii) Distance of Operation. The radar should be capable of displaying distinct and identifiable targets throughout the angular range of the display and at approximately 80 percent of the maximum range.
(iii) Beam Tilting. The radar antenna should be installed so that its beam is adjustable to any position between 10 degrees above and below the plane of rotation of the antenna.
Tilt calibration should be verified.
(iv) Contour Display (ISO Echo).
(A) If heavy cloud formations or rainstorms are reported within a reasonable distance from the test base, select the contour display mode. The radar should differentiate between heavy and light precipitation.
(B) In the absence of the above weather conditions, determine the effectiveness of the contour display function by switching from normal to contour display while observing large objects of varying brightness on the indicator. The brightest object should become the darkest when switching from normal to contour mode.
(v) Antenna Stabilization, When Installed. While in level flight at 10,000 feet or higher, adjust the tilt approximately 2-3 degrees above the point where ground return was eliminated. Roll right and left approximately 15 degrees, then pitch down approximately 10 degrees (or within design limits). No ground return should be present.
(vi) Ground Mapping. Fly over areas containing large, easily identifiable landmarks such as rivers, towns, islands, coastlines, and so forth. Compare the form of these objects on the indicator with their actual shape as visually observed from the cockpit.
(vii) Mutual Interference. Determine that no objectionable interference is present on the radar indicator from any electrical or radio/navigational equipment when operating and that the radar installation does not interfere with the operation of any of the airplane's radio/navigational systems.
(viii) Electromagnetic Compatibility. With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(ix) Light Conditions. The display should be evaluated during all lighting conditions, including night and direct sunlight.
(11) Area Navigation.
(i) Advisory Circular 90-96. This AC is the basic criteria for evaluating an area navigation system, including acceptable means of compliance to Title 14 of the Code of Federal Regulations (14 CFR).
(ii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(12) Inertial Navigation.
(i) Basic Criteria. Advisory Circular 25-4 is the basic criteria for the engineering evaluation of an inertial navigation system (INS) and offers acceptable means of compliance with the applicable 14 CFR Part. The engineering evaluation of an INS should also include an awareness of AC 121-13, which presents criteria to be met before an applicant can get operational approval. For flights up to 10 hours, the radial error should not exceed 2 n.m. per hour of operation on a 95 percent statistical basis. For flights longer than 10 hours, the error should not exceed ±20 n.m. cross-track or ±25 n.m.
along-track error. A 2 n.m. radial error is represented by circle, having a radius of 2 n.m., centered on the selected destination point.
(i) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(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. See AC 23.1311-1.
(16) LORAN C Navigation Systems. See AC 20-121A.
(17) Microwave Landing Systems. (RESERVED)
(18) Global Positioning System Navigation. (RESERVED)
288. RESERVED
289. SECTION 23.1303 FLIGHT AND NAVIGATION INSTRUMENTS.
a. Free Air Temperature (FAT). Section 23.1303(a)(4) requires that reciprocating enginepowered airplanes of more than 6,000 lbs. 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 302 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 137
of safety, derived from that reduction. Additional efforts would not result in any significant improvements in reliability.
290. SECTION 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. Advisory Circular (AC) 23-17 covers the installation of fuel flowmeters in airplanes with continuous-flow fuel injection reciprocating engines.
291. SECTION 23.1307 MISCELLANEOUS EQUIPMENT. (RESERVED)
292. SECTION 23.1309 EQUIPMENT, SYSTEMS, AND INSTALLATIONS. (RESERVED) This item is covered in AC 23.1309.
293.-299. RESERVED.
Section 2. INSTRUMENTS: INSTALLATION
300. SECTION 23.1311 ELECTRONIC DISPLAY INSTRUMENT SYSTEMS.
This item is covered in AC 23.1311-1.
301. SECTION 23.1321 ARRANGEMENT AND VISIBILITY. (RESERVED)
302. SECTION 23.1322 WARNING, CAUTION, AND ADVISORY LIGHTS. (RESERVED)
303. SECTION 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 indicated airspeed (IAS), calibrated airspeed (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 138
Publication 1046 (see subparagraph 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 pitotstatic pressures into electronic signals. Instrument errors are determined for inflight conditions in steady state conditions. Ground run system calibrations 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) and 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(c) requires an airspeed calibration for use during the acceleratetakeoff 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. See Appendix 9 for test procedures and a sample data reduction.
c. Trailing Bomb or Airspeed Boom Method. See Appendix 9 for procedures, test conditions, and a sample data reduction.
d. Pace Airplane Method. See Appendix 9 for test procedures.
e. Tower Fly-by. See 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 accelerate-takeoff 139
ground run and is used to determine IAS values for various V1 and VR speeds. See 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.
304. SECTION 23.1325 STATIC PRESSURE SYSTEM.
a. Definitions. Paragraph 302 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:
() .ft,V5.661V2.01dV08865.0dh C
5.22 CC ⎥⎦⎤⎢⎣⎡ σ⎥⎥⎦⎤⎢⎢⎣⎡ ⎥⎦⎤⎢⎣⎡+=
where Vc = calibrated airspeed, knots σ = ambient air density ratio dV c = 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 VSO 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 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.) = [ ][ ])TEST()TEST(dh σ
Where:
dh (TEST) = Difference in test pressure altitude and airplane pressure altitude with associated instrument errors removed.
σ (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.
305. SECTION 23.1326 PITOT HEAT INDICATION SYSTEMS. (RESERVED)
306. SECTION 23.1327 MAGNETIC DIRECTION INDICATOR. (RESERVED)
307. SECTION 23.1329 AUTOMATIC PILOT SYSTEM.
This subject is covered in AC 23-17.
308. SECTION 23.1331 INSTRUMENTS USING A POWER SUPPLY. (RESERVED)
309. SECTION 23.1335 FLIGHT DIRECTOR SYSTEMS. (RESERVED)
310. SECTION 23.1337 POWERPLANT INSTRUMENTS.
a. Explanation.
(1) Fuel Quantity Indicator. The indicator should be legible and easily read without excessive head movement. The calibration units and the scale graduations should be readily apparent. Units should be consistent with AFM procedures and performance data.
(2) Auxiliary Tanks. A fuel quantity indicator 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.
311.-318. RESERVED.
Section 3. ELECTRICAL SYSTEMS AND EQUIPMENT
319. SECTION 23.1351 GENERAL. (RESERVED)
320. SECTION 23.1353 STORAGE BATTERY DESIGN AND INSTALLATION.
a. Explanation. When ascertaining that the installed airplane battery capacity is adequate for compliance with § 23.1353(h), account should be taken of any services or equipment essential for the continued safe flight and landing of the particular airplane in accordance with the approved emergency procedures and in any approved conditions of operation. Account should also be taken of those services that cannot readily be shed. In order to ensure that services will function adequately for the prescribed period, the duration of battery supply should normally be based on a battery capacity of 72 percent of the nameplate rated capacity at the one hour rate. This figure takes into consideration the battery state of charge, the minimum capacity permitted during service life and the battery efficiency and is based on a battery capacity of 80 percent of the nameplate rated, at the one hour rate, and a 90 percent state of charge. Recognition time may depend on the kind of warning systems.
Section 23.1353 (h) is as follows:
In the event of a complete loss of the primary electrical power generating system, the battery must be capable of providing at least 30 minutes of electrical power to those loads that are essential to continued safe flight and landing. The 30-minute time period includes the time needed for the pilots to recognize the loss of generated power and take appropriate load shedding action.
The airplane characteristics and all systems needed for continued safe flight and landing for all intended operations need to be considered. Assuming operations will be under Part 135 IFR conditions; and only addressing the specific requirements under Part 91 and 135 the current type certificate application for Part 23, amendment 23-53, single-pilot, day/night/VFR/IFR, with approval for flight-into-known-icing conditions.
The 30-minute of electrical power should include the following equipment/instruments if they require electrical power:
• One airspeed and altimeter, with a heated pitot tube and heated static pressure source if needed to comply with section 23.1325.
• Instrument light for the magnetic compass.
• One navigational equipment appropriate to the ground facilities to be used.
• One communications systems.
• One gyroscopic pitch and bank indicator.
• One clock.
• Any powerplant parameter necessary for continued safe flight and landing or for engine restart
• Electronic Engine Controls, if installed.
b. Procedures. None.
321. SECTION 23.1357 CIRCUIT PROTECTIVE DEVICES. (RESERVED)
322. SECTION 23.1361 MASTER 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.
323. SECTION 23.1367 SWITCHES. (RESERVED)
324.-328. RESERVED.
Section 4. LIGHTS
329. SECTION 23.1381 INSTRUMENT LIGHTS. (RESERVED)
330. SECTION 23.1383 LANDING LIGHTS. (RESERVED)
331.-335. RESERVED.
Section 5. SAFETY EQUIPMENT
336. SECTION 23.1411 GENERAL. (RESERVED)
337. SECTION 23.1415 DITCHING EQUIPMENT. (RESERVED)
338. SECTION 23.1416 PNEUMATIC DEICER BOOT SYSTEM. See AC 23-17.
339. SECTION 23.1419 ICE PROTECTION. See AC 23-17 and AC 23.143-1.
340.-349. RESERVED.
Section 6. MISCELLANEOUS EQUIPMENT
350. SECTION 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.
351. SECTION 23.1435 HYDRAULIC SYSTEMS. (RESERVED)
352. SECTION 23.1441 OXYGEN EQUIPMENT AND SUPPLY. (RESERVED)
353. SECTION 23.1447 EQUIPMENT STANDARDS FOR OXYGEN DISPENSING UNITS.
(RESERVED)