402(a)(3)(D)(ix) Airframe noise [ETM 4.2.1.3.4.9] To account for the contribution of airframe noise, measured flight datum airframe noise on its own, or combined with an approved airframe noise analytical model, may be used to develop an airframe noi se database. The airframegenerated noise, which can be treated as a point source for adjustment purposes, is normalized to the same conditions as those of the other (i.e., engine) sources, with due account given for the effects of spherical divergence, atmospheric absorption and airspeed as described in Appendix A of part 36.
Airframe noise for a specific configuration varies with airspeed (see Reference T5) as follows:
Δܮܲܵ ൌ 50logሺܸோாி/்ܸாௌ்ሻ where:
— VREF is the approved reference airspeed for the “flight datum” airplane; and — V TEST is model or measured airspeed.
The above equation is also valid for adjustments to EPNL where an empirically derived coefficient replaces the coefficient 50 since that number may be somewhat configuration-dependent. However, the approval of the FAA is required for values other than 50.
402(a)(3)(D)(x) Airplane flight path considerations [ETM 4.2.1.3.4.10] When computing the one-third octave band sound pressure le vels corresponding to the slant distance of the airplane in flight from the noise measuring point, the principal effects are spherical divergence (inverse square law) adjustments from the nominal static distance and sound attenuation due to atmospheric absorption (as described in Appendix A of part 36). Furthermore, account should be taken of the difference between the static engine axis and that axis in flight relative to the reference noise measuring points. The adjustments should be applied to the component noise source levels that have been separately identified.
402(a)(3)(D)(xi) Total noise spectra [ETM 4.2.1.3.4.11] Both the engine tonal and broadband noise source components in flight, together with the airframe noise and any installation effects, are summed up on a mean-square pressure basis to construc t the spectra of total airplane noise levels.
During the merging of broadband and tonal components, consideration should be given to appropriate band-sharing of discrete frequency tones.
The effects of ground reflections must be included in the estimate of free-field sound pressure levels in order to simulate the sound pressure levels that would be measured by a microphone at a height of 4 ft (1.2 m) above a natural terrain. Information in Reference T4 or T15 may be used to apply adjustments to the free-field spectra to allow for flight measurements being made at 4 ft (1.2 m). Alternatively, the ground reflection adjustment can be derived from other approved analytical or empirically derived models. Note that the Doppler adjustment for a static source at frequency (fstatic) applies to a moving source (i.e., airplane) at a frequency (fflight) where ݂௧ ൌ݂ ௦௧௧/ሺ1 െ ߣݏܿܯሻ using the same terminology as described above for the Doppler effect. This process is repeated for each measurement angle and for each engine power setting. With regard to lateral attenuation, the information in Reference T14, applicable to the computation of lateral noise may be applied.
402(a)(3)(D)(xii) EPNL computations [ETM 4.2.1.3.4.12] For EPNL calculations, a time is associated with each extrapolated spectrum along the flight path. Note that the time is associated with each measurement location with respect to the engine/airplane reference point and the airplane’s true airspeed along the reference flight path, assumi ng zero wind. For each engine power setting and minimum distance, an EPNL is computed from the projected time history using the methods described in Appendix A of part
36.
402(a)(3)(D)(xiii) Chan ges to noise levels [ETM 4.2.1.3.4.13] An NPD plot can be constructed from the projected static data for both the original (i.e., flight datum) and the changed configurations of the engine or nacelle tested. Comparisons of the noise versus engine thrust (power) relationships for the two configurations at the same appropriate minimum distance will determine whether or not the changed configuration resulted in a change to the noise level from an engine noise source. If there is a change in the level of source noise, a new in-flight airplane NPD plot can be developed by adjusting the measured original NPD plot by the amount of change indicated by the comparison of the static-projected NPD plots for the original and changed versions within the limitations specified in 402(a)(3)(ii) for EPNL.
The certification noise levels for the derived version of an airplane may be determined from NPD plots at the relevant reference engine power and distance, with an additional adjustment of [10 log Vnom/Vr] for the velocity of the airplane at the certification reference condition relative to the nominal velocity (Vnom) used in developing the NPD plots.
402(b) Propeller-Driven Airplanes Over 19,000 lb [ETM 4.2.2] The procedures described in this chapter have been used as equivalent in stringency for propeller-driven airplanes with maximum certificated takeoff weight exceeding 19,000 lb, as provided in Appendix B.
402(b)(1) Flight test procedures [ETM 4.2.2.1] 402(b)(1)(A) Flight path intercept procedures [ETM 4.2.2.1.1] Flight path intercept procedures, as described in 402(a) (1)(i), have been used to meet the noise certification demonstration requirements in lieu of full takeoffs and/or landings.
402(b)(1)(B) Generalized flight test procedures [ETM 4.2.2.1.2] Generalized flight test procedures, other than normal noise demonstration takeoffs and approaches, have been used to meet two equivalency objectives:
a) NPD plots Noise data are acquired over a range of engine power settings at one or more heights. This information permits the development of generalized noise characteristics necessary for the certification of a “family” of similar airplanes. The procedures used are similar to those described in 402(a)(1)(ii), with the exception that the NPD plots employ engine noise performance parameters (µ) of propeller helical tip Mach number (MH) and shaft horsepower (SHP/δamb) (see Figure 34), where δamb is defined in 402(a)(1)(ii)(A).
Engine noise performance parameter ( ) SHP / amb Helical tip Mach number MH EPNL at a given height Figure 34. Form of noise-power-distance (NPD) plot for heavy propeller-driven airplanes In order to ensure that propeller inflow angles are similar throughout the development of the noise-sensitivity data as the airplane weight changes, the airspeed of the airplane used in the flight tests for developing the lateral and flyover data must be V2 + 10 kt (V2 + 19 km/h) to within ±3 kt (±5.5 km/h), as appropriate for the weight of the airplane during the test.
For the development of the NPD data for the approach case, the speed and approach angle constraints imposed in B36.7(c) and B36.8 of part 36 cannot be satisfied over the typical range of power needed. For the approach case, a steady speed of V REF + 10 kt (V REF + 19 km/h) should be maintained to within ±3 kt (±5.5 km/h) and the flyover height over the microphone should be 400 ± 100 ft (122 ± 30 m). Within these constraints, the test approach angle at the test power should be that which results from the test airplane conditions (i.e., weight, configuration, speed and power).
b) Noise level changes Comparisons are made of flyover noise test data for different developments of an airplane type (e.g., a change in propeller type). Such changes are used to establish certification noise levels of a newly derived version as described in 402(a)(1)(ii).
402(b)(1)(C) Determination of th e lateral certification noise level [ETM 4.2.2.1.3] For propeller-driven airplanes, B36.3(b) of part 36 specifies a full-power measurement point under the flight path as a replacement for the lateral measurement point. This section describes appropriate equiva lent procedures for those airplanes for which the two-microphone lateral measurement method was applicable.
Determination of the lateral certification noise level employing an alternative procedure using two microphone stations located symmetrically on either side of the takeoff flight path similar to that described in 402(a)(1)(iii) has been approved. However, when this procedure is used, matching data from both lateral microphones for each fly-by must be used for the lateral noise determination. Cases where data from only one microphone is available for a given fly- Page 183
by must be omitted from the determination. The following paragraphs describe the procedures for propeller-driven heavy airplanes:
a) the lateral EPNL from propeller-driven airplanes, wh en plotted against height opposite the measuring sites, can exhibit distinct asymmetry; the maximum EPNL on one side of the airplane is often at a different height and noise level from that measured on the other side;
b) in order to determine the average maximum lateral EPNL (i.e., the certification sideline noise level) it is therefore necessary to undertake a number of flights over a range of heights to define the noise versus height characteristics for each side of the airplane; a typical height range would cover between 100 ft (30 m) and 1800 ft (550 m) above a track at right angles to, and midway along, the line joining the two microphone stations; the intersection of the track with this line is defined as the reference point;
c) since experience has shown the maximum lateral noise level may often be near the lower end of this range, a minimum of six good sets of data, measured simultaneously from both sides of the flight track, should be obtained for a range of airplane heights as low as possible; in this case takeoffs may be necessary; however care should be taken to ensure that th e airspeed is stabilized to at least V 2 + 10 kt (V 2 + 9 km/h) over the 10 dB-down period;
d) the airplane climbs over the reference point using takeoff power, speeds and configuration as described in B36.7(b)(3) and B.36.7(b)(4) of part 36;
e) the lateral certification noise level is obtained by find ing the peak of the curve of noise level (EPNL) corrected to reference-day atmospheric absorption values and plotted against airplane height above the reference point (see Figure 35); this curve is described as a least-sq uares curve fit through the data points defined by the median values of each pair of matche d data measured on each side of th e track (i.e., the average of the two microphone measurements for a given airplane height); and
f) to ensure that the requirements of A36.5.4.2 of part 36 are met, the 90 percent c onfidence limits should be determined in accordance with paragraph 305.
402(b)(1)(D) Measurements at non-reference points [ETM 4.2.2.1.4] In some instances, test measurement points may differ from the reference measurement points as specified in Appendix B. Under these circumstances an applicant may request a pproval of data that have been adjusted from actual measurements to the reference conditions for reasons described in 402(a)(1)(v).
Noise measurements collected closer to the test airplane than at the certification reference points are particularly useful for adjusting propeller noise data because they are dominated by low frequency noise. The spectra roll off rapidly at higher frequencies and are often lost in the background noise at frequencies above 5000 Hz. Section 306(c), describes a procedure for background noise adjustment.
x x x x x x x x x x x x x x x x x x x x x x x x x x xx x x Noise levels from right side of track Noise levels from left side of track Noise certification level EPNL Aeroplane height Figure 35. Typical lateral noise data plot for heavy propeller-driven airplanes Non-reference measurement points may be used provided that measured data are adjusted to reference conditions in accordance with the requirements of A36.9 of part 36 and that the magnitude of the adjustments does not exceed the limits cited in B36.8(f) of part 36.
402(b)(2) Analytical procedures [ETM 4.2.2.2] Equivalent analytical procedures rely upon the available noise and performance data of an airplane type. The generalized relationships between noise levels, propeller helical tip Mach number, and shaft horsepower, as well as the adjustment procedures for speed and height changes in accordance with the methods of Appendix A are combined with certificated airplane performance data in order to determine noise level changes resulting from type design changes. The noise level changes are then added to or subtracted from the certifica tion noise levels that are demonstrated by flight test measurements for the “flight datum” airplane.
Certifications using analytical procedures have been approved for type design changes that result in predictable noise level differences. The type design changes include the following:
a) an increase or decrease in maximum takeoff and/or landing weight of the airplane from the originally certificated weight;
b) power increase or decrease for engines that are acoustically similar and fitted with propellers of the same type;
c) airplane engine and nacelle configuration changes, usually minor in nature, including derivative airplane models with changes in fuselage length and flap configuration; care is however needed to ensure that the Page 185
existing noise sources are not modified by these changes (e.g., by changing the flow field into the propellers);
and
d) minor airframe design changes that could indirectly affect noise levels because of an impact on airplane performance (e.g., increased drag ); changes in airplane performa nce characteristics derived from aerodynamic analysis or testing have been used to demonstrate how these changes can affect the airplane flight path and consequently the demonstrated noise levels of the airplane.
402(b)(3) Ground static testing procedures [ETM 4.2.2.3] 402(b)(3)(A) General [ETM 4.2.2.3.1] Unlike the case of a turbojet or turbofan powerplant, static tests involving changes to the propeller are not applicable for determining noise level changes in the development of a propeller-driven airplane /powerplant family because of changes in the aero-acoustic operating conditions of the propeller when run statically compared with conditions existing during flight. The propeller noise levels measured during a static test can include significant contributions from noise source components that are not normally important in flight. However, limited static tests on engines with propellers, which are used as engine loading devices, can be utilized to determine small noise changes, as described below.
402(b)(3)(B) Guidance on the test-site characteristics [ETM 4.2.2.3.2] Guidance on the test-site characteristics, data acquisition and analysis systems, microphone locations, acoustical calibration and measurement procedures for static testing is provided in Reference T13 and is equally valid in these respects for propeller powerplants (see 402(a)(3)(iii)).
402(b)(3)(C) Static tests of the gas generator [ETM 4.2.2.3.3] Static tests of the gas generator can be used to identify noise changes resulting from changes to the design of the gas generators, or to the internal structure of the engine, in the frequency ranges where:
a) there is a contribution to the airplane EPNL;
b) that part of the spectrum is clearly dominated by the gas generator; or
c) ancillary equipment under circumstances where th e propeller and its aerodynamic performance remains unchanged.
Such circumstances where the propeller and its aerodynami c performance remain unchanged include, for example, changes to the compressor, turbine or combustor of the powerplant. The effect of such changes should be conducted under the same test, measurement, data reduction and extrapolation procedures as described in 402(a)(3) for turbojet and turbofan engines. The noise emanating from any propeller or other power extraction device used in static tests should be eliminated or removed analy tically. For the purposes of airplane EPNL calculation, the measured “flight datum” airplane propeller contributions should be included in the computation process.
402(c) Helicopters [ETM 4.2.3] The objective of a noise certification test is to acquire da ta for establishing an accurate and reliable definition of a helicopter’s noise characteristics (see Appendix H of part 36) . This Appendix establishes a range of test conditions and procedures for adjusting measured data to reference conditions.
402(c)(1) Flight test procedures [ETM 4.2.3.1] 402(c)(1)(A) Helicopter test speed [ETM 4.2.3.1.1] There are two requirements for helicopter test speeds. Firstly, the airspeed during the 10 dB-down period should be close to the reference speed (i.e., within ±5 kt (±9 km/h), (see H36.103(b) of part 36) in order to minimize speed adjustments for the three certification conditions of takeoff, flyover and approach.
The second speed requirement applies to the flyover case (see H36.105(c) & (d) of part 36). The number of level flyovers made with a headwind component must be equal to the number of flyovers made with a tailwind component.
The objective is to minimize the effect of wind on the measured flyover no ise levels. If, however, the absolute wind speed component in the direction of flight , as measured at a height of 33 ft (1 0 m) above ground, is less than ±5 kt (±2.6 m/s), then the effect of wind direction can be considered to be negligible. In this case, the measured flyover can be considered to be either a headwind or tailwind test run.
The applicant may find that although there are at least three valid flyovers with a headwind component and three valid flyovers with a tailwind component, there are more valid flyovers with one wind component than with another. In this case, the applicant will need to discuss with the FAA which flyovers are to be used in the determination of the final EPNL value for flyover. In many cases, preference may be given to using level flyovers performed in pairs in order that the meteorological conditions are as similar as possible for the two flyovers in each pair. Hence, there is merit in considering conducting flyovers in pairs for all wind speed conditions. Each pair should consist of two flyovers performed one after the other in opposite directions along the reference flight track.
The measurement of ground speed may be obtained by timing the helicopter as it passes over two points at a known distance apart on the helicopter track during the flyover noi se measurements. These two points should straddle the noise measurement microphone array.
402(c)(1)(B) Atmospheric test conditions [ETM 4.2.3.1.2] The temperature, relative humidity and wind velocity limitations are contained in A36.2.2.2 of part 36. The parameters are measured at 33 ft (10 m) at a location subject to approval by FAA. For adjustment purposes the measured values of these parameters are assumed to be representative of the air mass between the helicopter and the microphones. No calculation procedures based on the division of the atmosphere into layers are required, but such a method of analysis could be accepted by the FAA.
402(c)(1)(C) Temperature and relative humidity measurements [ETM 4.2.3.1.3] Temperature and relative humidity measurements, as defined in H36.101(c) of part 36, have to be made at a height of 33 ft (10 m) above the ground. The measured values are used in the adjustment of the measured one-third octave band sound pressure levels for the effects of atmospheric absorption to account for the difference in the sound attenuation coefficients in the test and reference atmospheric conditions as given in H36.205(f) of part 36. The distances AL, ALr, AM, AMr, AN and ANr in the equations of H36.205(f) refer to the distances between positions on the measured and reference flight paths corresponding to the PNLTM position and the noise measurement point.
As a consequence the procedure assumes that the differen ce between the temperature and relative humidity at 33 ft (10 m) and the PNLTM position is zero, or small, and that the atmosphere can be represented by the values measured at 33 ft (10 m) above the ground in the vicinity of the noise measurement point Data obtained from European and U.S.
certification tests over a number of years and records provided by the United Kingdom meteorological office confirm that this assumption is valid over a wide range of meteorological conditions.
Noise certification measurements may be made under test conditions where significant changes in temperature and/or relative humidity with height are expected. Of particular concern are conditions when a significant drop in humidity with altitude is expected. Such special conditions might be encountered in desert areas shortly after sunrise where the temperature near the ground is lower and the relative humidity considerably higher than at the height associated with the PNLTM point. Measurements made under such conditions should be adjusted by using the average of the temperature and relative humidity measured at 33 ft (10 m) above the ground a nd at the height associated with the PNLTM point in order to eliminate errors associated with the use of data measured at 33 ft (10 m) only (see also 402(c)(1)(v)).
H36.101(c)(3) of part 36 limits testing to conditions wher e the sound attenuation rate in the 8 kHz one-third octave band is not more than 12 dB/100 m. If, however, the dew point and dry bulb temperature are measured with a device that is accurate to within ±0.5°C, it has been found acceptable by FAA to permit testing in conditions where the 8 kHz sound attenuation rate is not more than 14 dB/100 m.
402(c)(1)(D) Modifications or upgrades involving aerodynamic drag changes [ETM 4.2.3.1.4] The use of drag devices, such as drag plates mounted beneath or on the sides of the “flight datum” helicopter, has proven to be effective in the noise certification of modi fications or upgrades involving aerodynamic drag changes.
External modifications of this type are made by manufacturers and aircraft “modifiers”. Considerable cost savings are realized by not having to perform noise testing of numerous individual modifications to the same model series. Based on these findings it is considered acceptable to use the following as an equivalent procedure:
a) for helicopters to be certificated under Appendix H a drag device is used that produces the aerodynamic drag calculated for the highest drag modification or combination of modifications;
b) with the drag-producing device installed, an flyover test and, if considered appropriate by the FAA, takeoff and/or approach tests are performed by using the appropriate noise certification reference and test procedures;
c) a relationship of noise level versus change in aerodynamic drag or airspeed is developed by using noise data (adjusted as specified in Appendix H of part 36) of the “flight datum” helicopter and of the “high drag” configuration;
d) the actual airspeed of the modificati on to be certificated is determined from performance flight testing of the baseline helicopter with the modification installed; and
e) using the measured airspeed of the modification, certification noise levels are determined by interpolation of the relationship developed in item c).
Note.— Modifications or upgrades involving aerodynamic drag changes that do not require noise certification are described in 101(d)(4) of this AC.
402(c)(1)(E) Anomalous test conditions [ETM 4.2.3.1.5] H36.101(c)(5) of part 36 requires that the tests be conducted under conditions where no anomalous meteorological conditions exist. The presence of anomalous atmospheric conditions can be determined to a sufficient level of certainty by monitoring the outside air temperature (OAT) with the use of the aircraft instruments. Anomalous conditions which could impact the measured levels can be expected to exist when the OAT at 492 ft (150 m) is higher by 3.6°F (2°C) or more than the temperature measured at 33 ft (10 m) above ground level. This check can be made in level flight at a height of 492 ft (150 m) within 30 minutes of each noise measurement.
Since the actual heights associated with the PNLTM points will not be known until the analysis is made, measurements Page 188
of temperature and relative humidity can be made at a number of heights and the actual value determined from a chart of temperature and relative humidity versus height. Alternatively, since the influence of height is small, measurements at a fixed height in the order of 394 ft (120 m) and 492 ft (150 m) can be used depending on the flight condition and agreement of the FAA prior to the tests being conducted.
If tests are adjusted by using the “average” of the temperature and relative humidity measured at 33 ft (10 m) and the height associated with the PNLTM point (as described in the third paragraph of402(c)(1)(c)) then the provisions of the first paragraph of this section do not apply. The reason is that the impact of any anomalous meteorological conditions is taken into account by using the average of the temperature and relative humidity at 33 ft (10 m) and the height associated with the PNLTM point.
402(c)(1)(F) Helicopter test rotor speed [ETM 4.2.3.1.6] Operational rotor speed modes (e.g., CAT A) can form part of the normal procedures of the RFM and are used under specific operational circumstances. They typically involve airspeed ranges below those of the certification reference procedures. However, in some cases, such as a high pilot workload in the final approach phase, combined with instrument flight rules (IFR) conditions, their use has been permitted at higher airspeeds which includes the reference speed for noise certification. Hence, the maximum normal operating rotor speed corresponding to the reference flight condition should take into account any relevant operational rotor speed mode. The decision on how and which operational rotor speed modes are to be applied for noise cer tification is normally coordinated with the flight test experts of the FAA and is dealt with on a case-by-case basis.
402(c)(1)(G) Helicopter test weight [ETM 4.2.3.1.7] The weight of the helicopter during the noise certification demonstration (see H36.101(b)(6) & (8) of part 36) must lie within the range of 90 percent to 105 percent of the maximum takeoff weight for the takeoff and flyover demonstrations and between 90 percent to 105 percent of the maximum landing weight for the approach demonstration. For noise certification purposes the effect of change of weight is to change the test-day flight path for takeoff, and adjustments to the reference flight path should be made for spherical spreading and atmospheric attenuation as described in H36.205(f) of part 36.
In some cases, such as when the test aircraft weight is re stricted to a value somewhat less than the anticipated final certification weight, the applicant may, subject to the appr oval by the FAA, apply specific adjustments for weight variations. The applicant may be approved to use a 10-log relationship adjustment or otherwise determine, by flight test, the variation of EPNL with weigh. In such a case, the weight tested should include the maximum allowable test weight.
Note.— A similar adjustment procedure may be acceptable when the certificated weight is increased by a small amount subsequent to the flight tests.
402(c)(1)(H) Helicopter approach [ETM 4.2.3.1.8] H36.3(f)(ii) of part 36 constrains the approach demonstration to within ±0.5 of the reference approach angle of 6 .
Adjustments of the noise data to the reference approach angle are required to account for spherical spreading effects and atmospheric attenuation as described in H36.205(f) of part 36.
402(c)(2) Analytical procedures [ETM 4.2.3.2] 402(c)(2)(A) Helicopter test window for z ero adjustment for atmospheric attenuation [ETM 4.2.3.2.1] Page 189
There is currently a “test win dow” contained in H36.101(c) of part 36 which needs to be met before test results are acceptable to FAA. In addition if the test conditions fall within a “zero attenuation adjustment window” (see Figure
36), defined as the area enclosed by (2°C, 95 percent RH), (30°C, 95 percent RH), (30°C, 35 percent RH), (15°C, 50 percent RH) and (2°C, 90 percent RH), then the sound attenuation adjustment of the test data may be assumed to be zero.
(30°C, 95% RH) (2°C, 95% RH) (2°C, 90% RH) Relative humidity % Temperature °C 0 5 1 0 1 5 2 0 2 5 3 0 3 5 0 10 20 30 40 50 60 70 80 90 100 (15°C, 50% RH) (30°C, 35% RH) Figure 36. Appendix H zero attenuation adjustment window Accordingly the terms:
0.01 [ (i) – (i)0] AL, and
0.01 (i)0 (AL – ALr) from the equation for SPL(i)r H36.205(f) of part 36 become zero and the equation for SPL(i)r becomes:
SPL(i)r = 20 log (AL / ALr) Furthermore, in this equation AL and ALr may be replaced by the test and reference distances to the helicopter when the helicopter is over the center noise measuring point provided that all the measured points for a particular flight condition are:
a) flown in test conditions within the “zero attenuation adjustme nt window” defined in Figure 37;
b) for flyover, the height is 492 ± 30 ft (150 ± 9 m);
c) for approach, the height over the microphone is 394 ± 33 ft (120 ± 10 m)); and
d) for takeoff, the distance adjustment given in H36.111(c)(2) of part 36 is not greater than 2 EPNdB.
The total effect of both simplifications cited above is that the equation in H36.205(f) of part 36 becomes:
SPL(i)r = SPL(i) + 20 log (AF / ArFr) and the duration adjustment term specified in H.36.205(g) becomes:
2 = -7.5 log (AF / AFr) + 10 log (V / Vr) where AL is the measured distance from the helicopter to the noise measuring point when the helicopter is directly over the center noise measuring point and ALr is the reference distance.
402(c)(2)(B) Procedure for the determination of source noise adjustment [ETM 4.2.3.2.2] For demonstration of flyover reference certification noise levels, off-reference adjustments are normally made by using a sensitivity curve of PNLTM versus advancing blade tip Mach number deduced from flyovers carried out at different airspeeds around the reference airspeed. However, adjustment may be made by using an alternative parameter or parameters approved by the FAA. If the test aircraft is unable to attain the reference value of the advancing blade tip Mach number or the agreed reference noise correlating parameter, then an extrapolation of the sensitivity curve is permitted, provided that the data cover a range of values of the noise correlating parameter between test and reference conditions as agreed by the FAA. The advancing blade tip Mach number, or agreed noise correlating parameter, must be computed from as-measured data using true airspeed, on-board outside air temperature (OAT) and rotor speed. A separate curve of source noise versus advancing blade tip Mach number, or another agreed noise correlating parameter, must be derived for each of the three noise certification meas urement points (i.e., centerline, left sideline and right sideline). Left and right sidelines are defined relative to th e direction of the flight for each run. PNLTM adjustments are to be applied to each microphone datum using the appropriate PNLTM function.
In order to eliminate the need for a se parate source noise adjustment to the flyover test results, the following test procedure is considered acceptable when the correlating parameter is the main rotor advancing blade tip Mach number (Mr).
Each flyover noise test must be conducted such that:
a) the adjusted reference true airspeed (V ar) is the reference airspeed (Vr) specified in H36.105(c)(5) of part 36 adjusted as necessary to produce the same main rotor advancing blade tip Mach number as associated with reference conditions;
Note.— The reference advancing blade tip Mach number (Mr) is defined as the ratio of the arithmetic sum of the main rotor blade tip rotational speed (VT) and the helicopter reference speed (Vr) divided by the speed of sound (cr) at 77°F (25°C) 1135.6T/s (346.1 m/s) such that:
ܯ ൌ ்ܸ ܸܿ Page 191
and the adjusted reference true airspeed (Var) is calculated from:
ܸ ൌܿ ሺ శ ೝ ೝሻെ ்ܸ, where c is the speed of sound calculated from the on-board measurement of outside air temperature (see 308 of this AC).
b) the test true airspeed (V T) must not vary from the adjusted reference true airspeed (V ar) by more than 3 kt (±5.5 km/h) or an equivalent approved variation from the reference main rotor advancing blade tip Mach number (Mr);
c) in practice, the tests will be flown to an IAS which is the adjusted reference true airspeed (Var) corrected for compressibility effects and instrument position errors; and
d) the on-board outside static air temp erature must be measured at the flyover height just prior to each flyover.
Note 1.— The calculation of noise levels, including the adjustments, is the same as that described in Appendix H except that the need for source noise adjustment is e liminated. It should be emphasized that in the determination of the duration adjustment (2), the speed component of the duration adjustment is calculated as 10 log(VG/VGr) where VG is the test ground speed and VGr is the reference ground speed.
Note 2.— The symbol VG is denoted as the symbol V and VGr as Vr in Appendix H 403 TECHNICAL PROCEDURES INFORMATION [ETM 4.3]
403(a) Jet and Propeller-Driv en Airplanes and Helicopters [ETM 4.3.1]
403(a)(1) Adjustment of measured noise data to reference conditions [ETM 4.3.1.1] Sect ions A36.9.3 and A36.9.4 of part 36 provides for the use of either simplified or integrated methods for adjusting measured noise data to reference cond itions. Criteria for selecting the adjustme nt method are given in A36.9.1.2 of part 36. Guidance in the form of technical procedures for performing each of these two adjustment methods is provided in the following sections.
403(a)(2) Determination of reference-condition noise geometry [ETM 4.3.1.2] For either the simplified or integrated adjustment method, aircraft noise geometry must be determined in order to locate the aircraft position on the reference flight path.
The method described in this section illustrates the fundamental principles and key elements of aircraft noise geometry.
It is not recommended as the only acceptable method. Other methods may be preferable depending on the techniques used for acquisition and adjustment of test data. Acceptab le methods will though have within them key elements consistent with these principles. Note that all methods, including the method described in this section, and the manner in which they are implemented by the applicant, are subject to approval by the certificating authority.
The methodology presented is dependent on obtaining an aver age, straight-line flight path that represents the test aircraft position during noise measurements. This method is based on characterizing this average straight-line flight path by a set of single-point descriptors, which can be easily obtained from any method of aircraft TSPI measurement data. Geometry relative to each centerline (flyover and approach) and lateral microphone of interest is then determined Page 192