ܪோ ൌ ሺ8203 െ ܦହሻ ∗t a n ሾ s i nିଵሺ ோ ሻሿ 50,ଵଵ.ସ where:
— D50 is the sea level, ISA takeoff distance in feet to a height of 50 ft at the maximum certificated takeoff weight and maximum certificated takeoff power;
— RC is the sea level, ISA best rate of climb (ft/s) at the maximum certificated takeoff weight and the maximum power and engine speed that can be continuously delivered by the engine(s) during this second phase; and — Vy is the speed (kt) for the best rate of climb.
The performance data in many flight manuals are often presented in terms of SI units. Typically the takeoff distance, expressed in meters, is given to a height of 15 m and the rate of climb and airspeed are expressed in meters per second (m/s). In such instances, the expression for reference flyover height, HR in meters becomes:
ܪோ ൌ ሺ2500 െ ܦଵହሻ ∗t a n ሾ s i nିଵሺோሻሿ 15, where:
— D15 is the sea level, ISA takeoff distance in meters to a height of 15 m at the maximum certificated takeoff weight and maximum certificated takeoff power;
— RC is the sea level, ISA best rate of climb (m/s) at the maximum certificated takeoff weight and the maximum power and engine speed that can be continuously delivered by the engine(s) during this second phase; and — Vy is the speed (m/s) for the best rate of climb.
The performance figures can normally be found in the performance section of an AFM or pilot’s handbook. Note that for certain categories of aircraft, a safety factor may be applied to the takeoff and climb performance parameters presented in the flight manual. In the case of multi-engine aircraft, it may be assumed that one engine is inoperative during part of Phase 1 and during Phase 2. For the purpos e of calculating the “acoustic” reference flight path, the takeoff distance and rate of climb should be determined for all engines operating by using gross (i.e., without the factor applied) data.
In addition, the best rate of climb speed, Vy, used in the equation for H is defined as the true airspeed (TAS). However in the flight manual, speed is normally presented in terms of IAS. This should be corrected to the calibrated airspeed (CAS) by applying the relevant position error and instrument corrections for the airspeed indicator. These corrections can also be found in the flight manual. For an ISA day at sea level, the TAS is then equal to the CAS.
503(a)(4) Reference conditions for source noise adjustments [ETM 5.3.1.4] Sections G36.201(c) and G36.201(d) of part 36 describe how corrections for differences in source noise between test and reference conditions must be made.
The reference helical tip Mach number and engine power are defined for the reference conditions above the measurement point (i.e., the reference atmospheric conditions at the reference height, HR) The reference temperature at the reference height (H R) is calculated under ISA conditions (i.e., for an ambient sea level temperature of 59°F (15°C) and assuming a standard temperature lapse rate of 3.566°F (1.98°C) per 1000 ft).
The reference temperature, TR°C, can be defined as:
ܶோ ൌ 15 െ 1.98ሺܪோ/1000ሻ The reference atmospheric pressure, PR hPa, is similarly calculated at the reference height (HR) for a standard sea level pressure of 1013.25 hPa, assuming a standard pressure lapse rate such that:
ܲோ ൌ 1013.25ሾ1 െ ሺ6.7862 ∗ 10ିܪோሻሿହଷଶହ .
503(a)(5) Worked exampl e for the calculation of reference flyover height and the associated referen ce atmospheric conditions [ETM 5.3.1.5] 503(a)(5)(A) For reference flyover height calculation [ETM 5.3.5.1] In Table 15 extracts are presented from the performance section of a flight manual for a typical light, single-engine propeller-driven airplane.
The introduction contains a statement to the effect that the information is derived from “measured flight test data” and includes “no additional factors”.
The sea level, ISA takeoff distance in feet to a height of 50 ft at the reference conditions cited in Appendix G of part 36 can be read from the table of takeoff distances presen ted for a paved runway at the maximum certificated takeoff weight of 1920 lb. Thus D50 is 1370 ft.
The rate of climb (RC) at the reference conditions can similarly be read from the RC table. Thus RC is 1000 ft/min.
The climb speed associated with the RC figures is given as 80 knots indicated airspeed (KIAS). The corresponding true airspeed at the reference conditions cited in Appendix G of part 36 is equal to the IAS, corrected according to the airspeed calibration table at the appropriate flap setting of 0°. Thus Vy is 81 knots true airspeed (KTAS).
Entering these parameters into the equation for reference height expressed in feet (HR ft) given in 503(a)(3) gives:
ܪோ ൌ ሺ8203 1370ሻ ∗t a n|݊݅ݏିଵሺ1000/101.4 ∗ 81ሻ| 5 0 and so HR = 888 ft.
Table 15. Example of flight manual performance section SECTIONS. PERFORMANCE
1. INTRODUCTION The data processed in this section enable flight planning to be carried out for flights between airfields with various altitudes, temperatures and field lengths. The information is derived from measured flight test data using CAA-approved methoda and factors to cover all the conditions shown. The data assume average plot skill and an aucraft engine and propeller in good condition.
No additional factors are included, and it is the pilot's responsibility to apply safety factors which must notbe less than those ...
б. AIRSPEED CALIBRATION KIAS 60 70 80 90 100 110 120 130 180 flap KCAS 61 71 81 91 101 111 121 131 181 15ª KIAS 50 60 70 80 85 flap KÇAS 51 61 71 81 86 35° KIAS 50 60 70 80 85 flap KCAS 50 59 69 79 84 TAKEOFF DISTANCE PAVED RUNWAY (I) — CONDITIONS Flaps - 15° Rotation speed - 53 KIAS Weight - 1920 Ib Speed at 50 ft - 65 KIAS Power - Full throttle IS A 20°C IS A - 10°C ISA ISA + 10°C IS A + 20°C IS A + 30°C Airfield height Ground Total Ground Total Ground Total Ground Total Ground Total Ground Total (‡ł) roll to 50 ft roll to 50 ft roll to 50 ft roll to 50 ft roll to 50 ft roll to 50 ft Sea level 530 1230 565 1290 600 1370 700 1580 750 1715 840 1900 5000 1045 2835 1065 2435 1090 2580 1170 2670 1295 2840 1290 2905 10,000 10,465 3335 1490 3390 1510 3435 1575 3560 1610 3695 1670 3790
RATE OF CLIMB CONDITIONS Flaps up Full throttle Weight - 1920 I Speed - - 80 KLAS Pressure Rate of climb (ft'min) altitude (ft) IS A - 20°C IS A ISA + 10°C IS A + 20°C Sea level 1035 1000 915 825 1000 980 945 860 770 2000 925 890 805 720 3000 870 830 750 665 4000 815 775 695 610 5000 765 720 640 560 6000 700 665 585 505 7000 635 605 56Ũ 450 8000 570 550 475 395 9000 495 48L 410 10,000 415 405 335 KLA5 = hots indicated airspeed KCAS = knots cabbrated airspeed 503(a)(5)(B) For calculation of reference atmospheric conditions [ETM 5.3.1.5.2]
a) The reference temperature at the reference height, HR, is given by the equation for TR in 503(a)(4):
800 TR = 15 - 1.98 1000 and so TR = 13.24°C.
b) The reference pressure at the reference height is given by the equation for P R in 503(a)(4):
ܲோ ൌ 1013.25ሾ1 െ ሺ6.7862 ∗ 10ି ∗ 888ሻሿହଷଶହ and so PR = 981 hPa.
Chapter 6 GUIDELINES FOR HELICOPTERS NOT EXCEEDING 7000 LB EVALUATED UNDER APPENDIX J OF 14 CFR PART 36
CHAPTER TABLE OF CONTENTS
601 EXPLANATORY INFORMATION237
601(a) General237
601(b) Noise Measurements238
601(c) Noise Certification Test and Measurement Conditions238
601(c)(1) Atmospheric test conditions238
601(c)(2) Anomalous meteorological conditions239
601(c)(3) Winds speed239
601(c)(4) Anomalous wind conditions239
601(c)(5) Flight test conditions239
601(c)(6) Flight test conditions241
601(c)(7) Flight test conditions242
601(c)(8) Flight test conditions242
601(c)(9) Flight test conditions243
601(d) Noise Unit Definition243
601(d)(1) Units244
601(e) Measurement of Helicopter Noise Received on the Ground244
601(e)(1) Noise measurement system244
601(e)(2) Noise measurement procedures244
601(f) Adjustment of Test Results244
601(f)(1) Data adjustments244
601(g) Reporting of Data to the FAA244
601(g)(1) Data reporting245
602 EQUIVALENT PROCEDURES INFORMATION245
602(a) General245
602(b) Procedures for the Determination of Changes in Noise Levels245
602(b)(1) Modifications or upgrades involving aerodynamic drag changes245
602(b)(2) Testing of light helicopters outside Appendix J temperature and humidity limits246
601 EXPLANATORY INFORMATION [ETM 6.1] 601(a) General [ETM 6.1.1] Helicopters not exceeding 7000 lb (3175 kg) can be certificated under either Appendix H or J of the part 36. Helicopters exceeding 7000 lb (3175 kg) can be certificated only under Appendix H of part 36. Guidelines for helicopters certificated using Appendix H of part 36 are provided in Chapter 4 of this AC.
Unlike Appendix H of part 36 which requires takeoff, flyover and approach tests with noise measurements being made at three measuring points, compliance with Appendix J of part 36 is based on flyover tests only, with the noise data being obtained only at one microphone locat ed under the flight track. Flight pa th adjustments are simplified and the final results determined in terms of SEL instead of EPNL.
Also since the Appendix J procedure is based on flyover tests only, there are no trade-off provisions between flight conditions as allowed in Appendix H of part 36. However, if a helicopter not exceeding 3175 kg (7000 lb) fails to comply with the noise limit of Appendix J of part 36, certification of the helicopter under the Appendix H procedures is allowed.
601(b) Noise Measurements [ETM 6.1.2] See 301 of this AC for technical procedures generally app licable for noise certification tests of all aircraft types including those evaluated under the provisions of Appendix J of part 36. Procedures specific to Appendix J are presented in the following section.
601(c) Noise Certification Test and Measurement Conditions [ETM 6.1.3] 601(c)(1) Atmospheric test conditions AMC §J36.101(c)(2), (4), and (6) [ETM AMC No. 2 A4 2.2.2]
(1) Temperature/relative humidity test window Tests are permitted over the range of temperature and relative humidity specified in J36.101(c)(2) of part 36 and shown in Figure 49.
Reference condition (25°C; 70%) 10 dB/100 m Temperature (°C) Relative humidity (%) Test window Figure 49. Appendix J temperatu re/relative humidity test window
(2) Testing outside the temperature and relative humidity window If the limits of J36.101(c)(2) of part 36 cannot be met, but the tests can be conducted within the temperature/relative humidity limits specified in H36.101(c) of part 36, then the applicant may alternatively elect to use the equivalent procedure defined in 602(b)(2).
(3) Meteorological measurements Measurements of the meteorological conditions are required to be made using equipment approved by the FAA.
The temperature, relative humidity and wi nd measurements are required to be made in the vicinity of the noise measurement point at a height between 4 ft (1.2 m) and 33 ft (10 m). This allows the use of hand-held equipment but does not preclude the use of more complex measuring systems. Temperature and relative humidity may be measured by a hand-held psychrometer. This device measures the wet and dry bulb temperatures from which the relative humidity is obtained. Similarly wind measurements may be made using a hand-held device if its specifications comply with the provisions of J36.101(c) of part 36.
(4) Temperature and relative humidity Measurements of temperature and relative humidity should be made at intervals of not more than one hour to ensure that the test conditions remain in the required limits. It is advisable to make measurements for each flight in case it is required at a later time to verify the test conditions.
Section 6.2.2.2 provides an equivalent procedure to the specifications for temperature and humidity measurement given in J36.101(c)(2) of part 36.
601(c)(2) Anomalous meteorological conditions AMC §J36.101(c)(5) [ETM AMC No. 3 A4 2.2.2]
(1) OAT differential The presence of anomalous meteorological conditions can be reasonably determined by monitoring the outside air temperature (OAT) using the helicopter on-board temperature gauge. Anomalous conditions that could impact the measured levels can be expected to exist when the OAT at 492 ft (150 m) is higher than the temperature measured at a height between 4 ft (1.2 m) and 33 ft (10 m) above the ground by more than 2°C (3.6°F). This check should be made in level flight at a height of 492 ft (150 m) within 30 minutes of each noise measurement.
601(c)(3) Wind speed AMC §J36.101(c)(3) and (4) [ETM AMC No. 4 A4 2.2.2]
(1) Wind speed limitations Wind speed measurement points and limits are given in J36.101(c) of part 36. Wind speed measurement system specifications are given in J36.101(c)(3) of part 36. Measurements should be taken frequently and, if near the limit, at least prior to each flight to confirm that the requirements are met. Particular attention should be given to the crosswind component since this can often be the limiting factor during testing. If feasible, the reference flight path direction can be changed to reduce the impact of this requirement. These wind speeds should be recorded and included in the report of the noise certification program. Wind limits are intended to minimize the adverse effects of wind on helicopter noise generation and sound propagation.
601(c)(4) Anomalous wind conditions AMC §J36.101(c)(5) [ETM AMC No. 5 A4 2.2.2]
(1) Identification of anomalous winds Anomalous winds are difficult to quantify but, providing the helicopter can be easily flown within the flight path and airspeed limits defined in part 36, there is no excessive side slip or yawing of the helicopter and no indication of rough air, then the flights can be considered acceptable. In the case where wind effects are anticipated to be a likely problem, an agreement between the applicant and the FAA or designated observer should be reached prior to testing to determine the acceptability criteria. Normally such issues arise only with gusty wind conditions near the 10 kt (5.1 m/s) wind speed limit, high crosswind conditions or the presence of strong thermals.
601(c)(5) Flight test conditions GM §J36.105(b) [ETM GM A4 2.4]
(1) Flyover height A reference flyover height of 492 ± 50 ft (150 ± 15 m) above the ground at the noise measurement point is specified as indicated in Figure 50. The measured noise data must be adjusted for the effects of spherical spreading Page 239
Test Window between the helicopter test flight path and the reference flight path, and between reference airspeed and adjusted reference airspeed as specified in J36.205(b) of part 36.
(2) Flight path measurement The helicopter flight path is also required to be within ±10 of the vertical above the noise measurement point (see Figure 50). This requirement, along with the height test window, means that the helicopter has to fly through a height/off-track test window located directly above the noise measurement point. There is no requirement to determine the magnitude of the off-track distance, but it is necessary to show that the helicopter is within the required height and angular limits. The applicant may ther efore find merit in recording the off-track values for subsequent confirmation of compliance.
(3) Flight track: markings The helicopter has to fly on a straight path and be within ±10 of the vertical overhead the noise measurement point as shown in Figure 50. In order for this to be successfully accomplished the applicant should consider marking on the ground, in a manner that can be readily seen from the helicopter, the intended track and associated lateral limits. Brightly colored or day-glow markers or lights to mark the flight track are advisable. These markings will be very important in the case of a small helicopter where the on-board equipment may be the minimum required to comply with the airworthiness certification.
Reference height 150 m (492 ft) Flight track noise measurement station 15 m (50 ft) 10° 10° Test window Figure 50. Flight boundaries for flyover test condition
(4) Number of test runs A minimum of six flyovers, with an equal number of flights with headwind and tailwind components over the noise measurement point, are required. These test runs should be conducted in pairs, since the aim is to minimize the influence of wind speed and direction on the measur ed SEL. The tests in each pair should be conducted immediately one after another in order that the meteorol ogical conditions are as similar as possible for the two test runs. It should be possible to determine immediately after each test run if it meets the necessary requirements and thus relatively easy to establish when three pairs of valid test runs have been made. The applicant would also be advised to conduct one or two additional pairs of test runs to ensure that after all the test parameters have been examined a minimum of three valid pairs of test run results are available. If additional valid pair(s) of test runs are obtained these will be required to be included in the analysis to determine the arithmetic average SEL.
(5) Landing gear position If the helicopter has a retractable landing gear, the landing gear position for noise tests needs to be that used for Page 240
the cruise configuration.
601(c)(6) Flight test conditions AMC §J36.105(c), (d) and (e) [ETM AMC A4 2.4]
(1) Test period The test conditions have to be maintained, or held constant, over an adequate distance (time interval) to encompass the 10 dB-down period. The maximum A-weighted sound pressure level, dB(A) or L Amax, will normally occur when the helicopter is at, or just prior to, the position directly above the noise measuring point. Pre-test flights should be conducted to determine the 10 dB-down period and ensure this period is adequately captured by the noise measurement system. It is advisable that the helicopter flight test conditions are stabilized well in advance of the initial 10 dB-down point and maintained until well after the second 10 dB-down point to ensure a valid noise measurement is obtained.
(2) Maximum normal operating rotor speed In order that the noise levels are representative of normal operation, the rotor speed used must be the maximum normal value associated with flyover at the reference conditions. Also, since on most helicopters small changes in rotor speed occur during a stabilized flight, a ±1 percent rpm variation is allowed.
(3) Test weight Fuel, together with the weight of the pilot, flight observers and ballast, are normally used to set the weight of the test helicopter within the required test range of +5 percent/–10 percent of the maximum takeoff weight. Fuel burn
(i.e., decrease in fuel weight) should be documented to determine the actual test value. Care must be taken regarding the location of ballast to ensure it does not have any adverse impact on the applicable center of gravity limits.
Unlike in Appendix H, there is no requirement to conduct any test run above the maximum takeoff weight. Note that variation of the flyover noise levels within the a llowable weight limits is small and for this reason no adjustments for difference in test weights are required. The helicopter weight, or the quantity of fuel from which the weight of the helicopter can be calculated, should be recorded for each flight.
Note.— Conducting the tests as near as possible to the upper weight limit of +5 percent can also be useful in supporting noise certification of future increases in maximum gross weight of the helicopter by minimizing the likelihood of having to conduct new tests.
(4) Background noise Some initial pre-test flyovers should be performed to confirm helicopter noise levels exceed background noise by 15 dB(A) as specified in J36.109(f)(3) of part 36. The FAA has generally accepted that the requirement has been met if the maximum helicopter noise level exceeds backgr ound noise levels by 15 dB(A). If this requirement cannot be met when the flyover test is conducted at 492 ft (150 m), a lower flyover height approved by the FAA may be used. This normally will be required only in the case of lightweight/small helicopters or those that generate extremely low noise levels.
Variations in measured noise levels of up to ±1.5 dB(A) from flight to flight may typically occur. The applicant should therefore ensure that the difference between background noise and helicopter noise levels is adequate for the quietest flyover noise measurements anticipated. Such information may also be useful for adjusting the sensitivity of the noise measurement system. The level of background noise may be influenced by the location of the test site.
601(c)(7) Flight test conditions GM §J36.105(c)(1) [ETM GM A4 2.4.2]
(1) Adjusted reference airspeed The flyover adjusted reference airspeed (Var) is defined as the value at the test temperature that gives the same advancing main rotor blade tip Mach number (MAT) as associated with the reference temperature of 25C (77ºF). On most helicopters the controlling noise source is dependent on the advancing blade tip Mach number. The advancing blade tip Mach number is dependent on the temperature and thus the sound level varies with temperature. To avoid the need to make a source noise correction, as would be required for the flyover tests under Appendix H of part 36 unless an equivalent procedure is used, tests to meet the requirements of Appendix J of part 36 are required to be conducted at an adjusted reference airspeed, Var, which gives the same advancing blade tip Mach number at the time of the test as would occur if the test were conducted under the reference conditions. The speed of sound increases with absolute temperature so that tests conducted in temperatures below the reference value of 25C (77ºF) at the reference flyover height will result in a higher advancing blade tip Mach number, and a reduction in test airspeed will be required to obtain the reference advancing blade tip Mach number. Similarly when the air temperature at flyover reference height is higher than 25 C (77ºF), the flyover test speed must be increased. This requires knowledge of the OAT measured on-board the aircraft at the time of the test. The applicant should note that it is essential to test at the required airspeed value since there is no provision for the adjustment of data obtained at the wrong Mach number.
601(c)(8) Flight test conditions AMC §J36.105(c)(1) [ETM AMC No. 1 A4 2.4.2]
(1) Test speed for light helicopters For the purposes of compliance with Appendix J of part 36, the helicopter should be flown at the test speed (Var) that will produce the same advancing blade Mach number (M R) as the reference speed at reference conditions given in J36.3(a) and J36.3(c) of part 36.
The reference advancing blade Mach number (MR) is defined as the ratio of the arithmetic sum of the main rotor blade tip rotational speed (VTIP) and the helicopter true airspeed (VREF) divided by the speed of sound (c) at 25°C
(346.1 m/s) such that:
ܯோ ൌ ܸோாி ܸ ்ூܿ The test airspeed (Var) is calculated from:
ܸ ൌܿ ்ሺܸோாி ܸ ்ூሻെܸ ்ூܿ where cT is the speed of sound obtained from the on-board measurements of outside air temperature.
Since the ground speed obtained from the flyover test s will differ from that for reference conditions, an adjustment, ΔJ3, of the form ΔJ3 = 10 log (Var/VREF) will need to be applied. ΔJ3 is the increment in decibels that must be added to the measured SEL.
There are two additional requirements for light helicopter test speed. First, the airspeed during the 10 dB-down period should be close (i.e., within ±3 kt (±5.5 km/h)) to the adjusted reference speed (see J36.105(c)(1)(ii) of part 36).
The second speed requirement states that the level flyovers must be made in equal numbers with a headwind component and tailwind component (see J36.105(b) of part 36). For practical reasons, if the absolute wind speed component in the direction of flight, as measured at a height between 4 ft (1.2 m) and 33 ft (10 m) above ground (see J36.101(c)(4) of part 36), is less than 5 kt (2.6 m/s), then the effect of wind can be considered to be negligible.
In this case, the measured flyover may be used to satisfy a test run in either the headwind or tailwind direction if the flyovers are conducted in pairs. Each pair should consist of two flyovers performed one after the other in opposite directions over the reference flight track.
Any changes in rotor speed, which may occur with the flight airspeed, will also need to be taken into account in the above calculations to determine the adjusted reference airspeed. If this is likely to occur then this topic should be reviewed with the FAA to determine if any additiona l adjustments to the flight test speed are required.
Normally this is not a concern, since the rotor speed will be independent of flight speed.
The applicant should also no te that the calculated adju sted reference airspeed (Var) is the adjusted true airspeed (TAS). Additional information will be required to determine the IAS for use by the pilot. This will normally be based on calibration charts or adjustments for the airspeed measurement system showing the IAS/TAS relationship.
601(c)(9) Flight test conditions AMC §J36.105(c)(2) [ETM AMC No. 2 A4 2.4.2]
(1) Rotor speed The rotor speed can be varied on some helicopters, and on others variations in the rotor speed can occur with flight speed. In order that the noise levels are representative of normal operation, the rotor speed used must be the maximum normal value associated with flyover at the reference conditions.
Note.— It is not the intent to require noise measurements at any value but the maximum used during normal operations, and thus testing at the maximum tolerance rotor speed is not required.
On some helicopters two distinct rotor speed values are available. If both can be used for normal operations then the noise certification has to be conducted at the higher roto r speed. If the higher of the two speeds is restricted to special operations, or if the helicopter is configured so that it cannot be used at the reference conditions and/or during lower altitude flight, then, subject to approval by the FAA, testing at the lower rotor speed may be allowed.
On some helicopters it may be possible for the rotor speed to be changed by pilot action. In these cases noise certification will require the highest rotor speed specified in the RFM for the flyover flight condition at the maximum takeoff weight to be used.
For most turbine-engine-powered helicopters the rotor speed is automatically linked by the engine control/ governor system to the flight condition. If this results in a different rotor speed at the adjusted reference airspeed to that associated with the reference airspeed, then additional adjustments may be required to ensure the correct advancing blade tip Mach number is used for the tests. If this situation is likely to occur, FAA approval of the rotor speed and/or adjusted reference airspeed to be used should be obtained.
On some recently designed helicopters the use of lower ro tor speeds has been certifi cated for operations at low altitudes and/or in cruise flight. Since these lower rotor speeds are defined in the RFM and, if higher rotor speed values cannot be used at the reference conditions, exce pt possibly in an emergency, noise certification is conducted at the certificated lower rotor speed, subject to approval by the FAA. If, however, the helicopter simply incorporates a two-speed or multi-speed system, and either can be selected by the pilot, then the highest value is required to be used for noise certification.
601(d) Noise Unit Definition [ETM 6.1.4] Page 243
601(d)(1) Units GM §J36.109(b) [ETM GM 3.1] The noise levels are to be determined in terms of the SE L metric. The SEL is the time-integrated A-weighted sound level over the 10 dB-down period. This metric takes into account both the duration and the level of the sound.
601(e) Measurement of Helicopter Noise Received on the Ground [ETM 6.1.5] 601(e)(1) Noise measurement system AMC §J36.109(c) [ETM AMC A4 4.3]
(1) System and calibration requirements The noise measurement system and system calibration requirements are specified in J36.109(c) through J36.109(e) of part 36 for compliance with Appendix J of part 36. The noise demonstration compliance test plan must include a description of the system to be used for the noise measurement. The FAA must approve the measurement system and calibration procedures in order to ensure that accurate measurements and results are obtained.
601(e)(2) Noise measurement procedures AMC §J36.109(e) [ETM AMC A4 4.4]
(1) Sound level integration period The A-weighted sound pressure level must be integrated over the 10 dB-down period. When using an integrating sound level meter where the start and stop times are selected manually, the actual test integration period should be slightly longer than the true 10 dB-down period. This will not have any significant impact on the SEL value, providing the integration period is only a few seconds longer, since the noise levels will be more than 10 dB(A) below the maximum sound level value.
601(f) Adjustment of Test Results [ETM 6.1.6] 601(f)(1) Data adjustments AMC §J36.205(b) and (c) [ETM AMC A4 5.0]
(1) Height adjustment In order to account for the differences between the test heights (H) and reference height of 492 ft (150 m), which influence both the spherical spreading of the noise and the duration of the 10 dB-down period, the ∆1 adjustment is applied.
(2) Airspeed adjustment In order to account for the differences between the adjusted reference airspeed (Var) and reference airspeed (VREF), which influence the duration of the 10 dB-down period, the ∆2 adjustment is applied. Variations in the ground speed, and hence duration, as a result of wind at the test height also occur, but since test runs are to be made with equal numbers with headwinds and tailwinds (see J36.105(b) of part 36), this effectively cancels out this effect and no additional adjustments for duration are required.
601(g) Reporting of Data to the FAA [ETM 6.1.7] Page 244