For variable pitch propellers, the source noise sensitivity curves are developed from data taken with the aircraft flying over a range of propeller speeds, typically three, at a fixe d torque or MAP in a manner similar to that described in 502(d)(1) where NMCP or NMAX would in this case be the maximum propeller speed at the maximum permitted torque or MAP. This is repeated for two lower torque or MAP values in order to establish a carpet plot of maximum A- weighted noise levels against propeller speed and torque, MAP or SHP.
A plot of maximum A-weighted noise level (LAmax), helical tip Mach number (MH) and torque or MAP is developed.
This plot is then used to derive the source noise adjustment (LAmax) which is the difference between reference and test conditions at the noise certification power.
Generally the test and reference engine SHP can be derived from the engine manufacturer’s performance curves.
However, where such curves are not available, a correction should be applied to the manufacturer’s published engine SHP, which is normally presented for a range of engine speeds under ISA and sea level conditions, in order to establish the engine power level under the test conditions of ambient temperature and air density. The correction is as follows:
a) for normally aspirated engines:
்ܲ ൌܲ ோ|ሺܶோ/்ܶሻ భ మ|ሾሺߪ െ 0.117ሻ/0.883ሿ ;and ’
b) for turbo-charged engines:
்ܲ ൌܲ ோ|ሺܶோ/்ܶሻ భ మ, where:
— PT and PR are the test and reference engine powers;
— TT and TR are the test and reference ambient temperatures; and — ߪis the air density ratio.
Note.— In this context reference denotes the reference conditions for which the engine SHP is known.
502(e) No-Acoustical Change Guidance for Derived Versions of Propeller-Driven Airplanes Certificated According to Appendix G [ETM 5.2.5] After the certification in their basic configuration, small propeller-driven airplanes are often modified, either by a TC change of the TC holder or by an STC from a supplier. These changes can be of a different nature such as weight increase/decrease, engine change, power change, propeller change, installation of vortex generators, fitting of winglets or external mounted equipment (cargo boxes, floats, etc.). With regard to noise and depending on their nature, some changes might have to demonstrate compliance with the applicable requirements by a new flight test, others by reevaluation of the original noise flights or by demonstrating a no acoustical change.
The propeller and the engine are the main noise sources of small propeller-driven airplanes. Parameters like the diameter, the number of blades, the rpm, the pitch, the blade tip shape or the geometry could have an impact on the propeller noise signature. As for the engine, noise signature could change by modifying the rotor assembly or the exhaust.
Several sections refer to the determination of a new reference height due to a change in performance data, here D50, Vy and RC. New performance data are accepted for the recalculation only if they are established by a method approved by the FAA.
The following sections are intended to provide guidance for applicants concerning an NAC demonstration.
502(e)(1) No-acoustical change guidance for airplanes fitted with fixed pitch propellers [ETM 5.2.5.1] (Reserved) Note.— This guidance is limited to variable pitch propellers only. According to G36.111(c)(2)(iv) of part 36, takeoff rpm must be maintained throughout the noise test runs. During climb at the best rate of climb speed, the fixed pitch propeller can generally not reach its maximum operating rpm value. Therefore the reference propeller speed is defined to be the average propeller rpm calculated from all valid runs. Changing the performance (e.g., due to an engine change or a change in weight) might change the average rpm adversely. Currently the amount of change in the propeller rpm for a fixed pitch propeller cannot be estimated analytically.
502(e)(2) No-acoustical change guidance for airplanes fitted with variable pitch propellers [ETM 5.2.5.2] 502(e)(2)(A) Engine change without power change [ETM 5.2.5.2.1] The engine is one of the main noise sources of the airplane. Changes to the engine can take many forms. They vary from small changes within one engine family, normally addressed by different characters within the engine designation, to a complete re-engine. In the latter case the compliance demonstration can in general be achieved only by new noise flight tests. In the former case it is often obvious that the change has no acoustical impact, and a simple statement should suffice to demonstrate that the change is not acoustically significant.
502(e)(2)(B) Power increase with out changing the propeller rpm [ETM 5.2.5.2.2] Increasing the power output of an engine without changing th e takeoff weight will increase the engine noise level at source but also improve the takeoff performance. A method subject to the approval by the FAA can be applied to evaluate the increase in engine noise source. This increase in engine noise source may be offset by the higher reference height. Takeoff distance is shortened, the climb rate is increased and therefore the flyover height over the microphone is increased. If it can be shown that this increase in reference flyover height offsets the increase in engine source noise, the change in engine power may be considered as an NAC.
An NAC is acceptable if it can be demonstrated that a minimum of six valid flights are within the “new” height window.
The effect of angle-of-attack changes may be included in the analysis by a method approved by the FAA. The method must be robust enough to account for the effects of performance and angle-of-attack changes on noise levels. If the analysis method shows that the noise level does not increase then the noise level of the unmodified airplane can be applied. Otherwise new testing should be required.
Note.— Section G36.111(c)(2)(iv) of part 36 defines that the microphone has to be passed at maximum takeoff power. If source noise sensitivity curves are established in accordance with the procedure laid down in 502(d) of this chapter the noise level can be adjusted up to the highest power covered by the sensitivity curve. In such a case the power correction determined by the sensitivity curve should be used instead of the general adjustment Δ3 = K3 log (PR/PT).
502(e)(2)(C) Change in weight [ETM 5.2.5.2.3] According to G36.111(a) of part 36, the flight tests must be initiated at the maximum takeoff weight. Only increases in Page 230
takeoff weight up to the maximum actually flown during the original flight tests can be accepted without new flight tests.
If it can be demonstrated that a further weight increase and the corresponding loss of performance do not adversely affect the noise level by more than 0.10 dB(A), the certificated noise level may be assigned to this weight without additional flight tests.
A change in the weight of the aircraft will lead to different performance characteristics. A new reference height with the new performance parameters has to be determined to demonstrate the influence on the noise level. Possible impact on the propeller speed should be taken into account.
Similar to the previous section, the effect of angle-of-attack changes may be included in the analysis by a method approved by the FAA.
502(e)(2)(D) Drag change [ETM 5.2.5.2.4] While a change in drag generally has no direct impact on the noise at source, it may have an indirect effect on noise level through a change in performance.
A drag change will in general be introduced by modifications such as the fitting of cargo pods or external fuel tanks, larger tires, floats, etc. In most cases, the change in ae rodynamic noise can be shown to be negligible for small propeller-driven airplanes. However, there may be cases where the aerodynamic noise generated by the modification has to be addressed. The drag change might change the performance characteristics of the aircraft D50, Vy and/or RC leading to a change in reference flyover height. The performance characteristics defined in the AFM are approved by the performance experts of the FAA. In some cases the pe rformance experts agree to ap ply the former performance parameters to the modified aircraft if the applicant can demonstrate that the performance is not worse than the one for the basic aircraft. Three different situations have to be considered:
a) the performance characteristics are bette r than those of the parent aircraft;
b) the performance characteristics are identical to those of the parent aircraft; and
c) the performance characteristics are wors e than those of the parent aircraft.
With regard to noise these three situations should be dealt with as follows:
a) in the case of situation a), independent of whether th e applicant decides to maintain the old performance data or to document the better performance in the AFM, an NAC can be granted and the noise level for the parent version can be applied to the modified aircraft;
b) in the case of situation b), the noise level for the parent version can be applied to the modified aircraft without further investigation; and
c) in the case of situation c), in general a new flight test is required.
502(e)(2)(E) Different blade count propeller [ETM 5.2.5.2.5] The effect of changing the number of the propeller blades on the noise level is difficult to determine by analytical procedures. Typically the applicant is obliged to perform a new flight test. Propeller noise prediction routines are highly sophisticated requiring extensive data sets which can in general be provided only by the propeller manufacturer.
The use of such propeller noise prediction routines has to be acceptable to the FAA to demonstrate an NAC.
502(e)(2)(F) Different blade tip shape [ETM 5.2.5.2.6] Page 231
503 In general rounded tips are quieter than squared ones. The change from squared to rounded blade tips can be accepted as an NAC if the rpm and diameter remain the same.
TECHNICAL PROCEDURES INFORMATION [ETM 5.3] 503(a) Worked Example of Ca lculation of Reference Flyover Height and Reference Conditions for Source Noise Adjustments (Appendix G) [ETM 5.3.1] 503(a)(1) Introduction [ETM 5.3.1.1] The reference flyover height for an airplane certificated to Appendix G of part 36 is defined at a point which is 8202 ft (2500 m) from the start-of-roll beneath a reference flight path determined according to the takeoff reference procedure described in G36.111(c) of part 36. An expression for the reference flyover height in terms of commonly approved performance data and an example of how such an expression may be worked are presented in this section. The relationship between the reference height and the conditions to which source noise corrections are to be made is also explained.
503(a)(2) Takeoff reference procedure [ETM 5.3.1.2] The takeoff reference procedure for an airplane certificated to Appendix G is defined in G36.111(c) of part 36 under sea level, ISA conditions, at maximum takeoff weight for which noise certification is requested. The procedure is described in two phases:
a) the first phase commences at “brakes release” and con tinues to the point where the aircraft reaches a height of 50 ft (15 m) above the runway; the point of interception of a vertical line passing through this point with a horizontal plane 50 ft (15 m) below is often referred to as “reference zero”; and
b) the second phase commences at th e end of the first phase and assume s the airplane is in normal climb configuration with landing gear up and flap setting normal for “second segment” climb.
Note.— The reference “acoustic” flight path ignores the “first segment” part of the flight path, during which the aircraft accelerates to normal climb speed and, where appropriate, landing gear and flaps are retracted.
503(a)(3) Expression for reference height [ETM 5.3.1.3] The reference flyover height is defined according to the takeoff reference flight path at a point 8202 ft (2500 m) from the start-of-roll for an airplane taking off from a paved, level runway under the following conditions:
a) sea level atmospheric pressure of 1013.25 hPa;
b) ambient air temperature of 59°C (15°C) (i.e., ISA);
c) relative humidity of 70 percent; and
d) zero wind.
This height can be defined in terms of the approved takeoff and climb performance figures for the conditions described above as follows:
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