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
for the test data, including sound emission coordinates (t, X, Y, Z) for each measured acoustic spectrum (k) in the acoustic spectral time-history data set. Once the sound emission coordinates have been identified, sound propagation distances and acoustic emission angles are calculated, which are used to determine the position of the aircraft on the reference flight path. For the integrated method of adjustment only, the series of positions on the reference flight path are then used to obtain the effective duration for each spectrum k for the reference condition acoustic data set.
403(a)(2)(A) Assumptions [ETM 4.3.1.2.1]
a) The test aircraft position during noise measurements can be represented by a straight-line flight path;
b) the ground reference system used in the figures, a right-handed coordinate system, is assumed to be fixed to the surface of a flat earth with th e X axis pointing along the reference ground track, the Y axis pointing to the left of the reference ground track, and the Z axis pointing up;
Note.— For airplanes the term “reference ground track” refers to the “extended centerline of the runway” or the “extended runway centerline” referred to in B36.3 of part 36.
c) the point on the ground directly beneath the centerlin e microphone is the origin of the XYZ coordinate system (X=0, Y=0, Z=0);
d) the X-coordinate value increases with time as the aircraft moves through the noise measurement test site;
e) the single-point flight path descriptors represent average values over the noise duration as defined in §36.9.4.2(b) of part 36; and
f) angular quantities are expressed in radians except where otherwise noted.
403(a)(2)(B) Steps involved [ETM 4.3.1.2.2]
a) Characterization of a straight-line average test flight path based on descriptors for a single point (see 403(a)(2)(C));
b) determination of test ai rcraft position at time of sound emission of each acoustic spectrum (see 403(a)(2)(D));
c) calculation of the geometric minimum distance between the test flight path and the microphone (see 403(a)(2)(E));
d) determination of test aircraft noise geometry (sound propagation distance, acoustic emission angle) for each acoustic spectrum (see 403(a)(2)(F));
e) determination of reference flight path (see 403(a)(2)(G));
f) determination of reference sound propagation dist ance for each acoustic spectrum (see 403(a)(2)(H)); and
g) determination of effective duration for each acoustic spectrum for the integrated procedure (see 403(a)(2)(I)).
403(a)(2)(C) Characterization of a straight-line average flight path based on descriptors for a single point [ETM 4.3.1.2.3] A straight-line flight path can be defined knowing the aircraft position, speed, and three-dimensional direction (vector) at a single point in time (see Figure 37). For the method described in this section the “single-point” descriptors are:
— tOH is the time at overhead (the time when the aircraft X coordinate = 0.0);
— XOH is the X coordinate of the centre line microphone, X = 0.0;
— YOH is the lateral offset of the aircraft from the reference ground track at tOH ;
— ZOH is the aircraft height above the reference X-Y ground plane at tOH;
— VG is the average groundspeed over the noise duration;
— γ is the average climb/descent angle; and — χ is the average lateral cross-track angle.
Note.— The average ground speed, VG, used in calculations for aircraft noise geometry is independent from any cockpit instrumentation, and is to be determined from the aircraft position measurements. For reference conditions, the reference groundspeed, VGR, is to be determined from the reference climb/descent angle and its relationship to the reference true airspeed value, VR.
Generate the straight-line average aircraft flight path t, X, Y, Z position time history from the single-point flight path descriptors at an appropriate sample rate (typically two times per second):
For any relative time, t(p):
X(p) = (t(p) – tOH) (VG cos (χ)) + XOH Y(p) = (t(p) – tOH) (VG sin (χ)) + YOH Z(p) = (t(p) – tOH) (VG tan (γ)) + ZOH Page 194
Figure 37. Single-point flight path descriptors 403(a)(2)(D) Determination of test aircraft position at time of sound emission of each acoustic spectrum [ETM 4.3.1.2.4] Sound emitted from the aircraft takes a finite time to propagate prior to being received at the measurement microphone.
During this time, the aircraft has travelled a finite distance along the flight path. Therefore, it is necessary to determine the time and position coordinates of the aircraft for the point of sound emission for each acoustic record k (see Figure
38).
The spectral time history of measured aircraft noise data includes a series of slow sample times as specified in A36.3.7.5 of part 36. For each of these measurement times, tm(k), the associated time of sound emission, tE(k), as well as the sound emission coordinates, X E(k), YE(k), ZE(k), can be determined using information about the microphone, the measured aircraft position and time, and the test speed of sound (c).
The microphone position descriptors are:
— XMIC is the longitudinal distance along the reference ground track, between microphone location and coordinate system origin (typically 0.0);
— YMIC is the lateral distance between microphone location and the reference ground track (typically 0.0 for centerline microphone, and ±492 ft (±150 m) or ±1476 ft (±450 m) for lateral microphones);
— ZMIC is the height of ground at microphone location relative to the reference ground plane (typically 0.0);
and — HMIC is the height of microphone above local ground (4 ft (1.2 m));
Figure 38. Aircraft position at time of sound emission Using the average straight-line flight path position time history from 403(a)(2)(C) and the microphone position descriptors listed above, develop an emission/reception arra y for eventual determination of aircraft sound emission coordinates for each measured acoustic spectral data record, k, as follows:
a) Calculate sound speed (c) for test-day conditions (see 308);
b) For each pth aircraft position sample in the position time history, t( p), X(p), Y(p) and Z(p ), calculate the following:
1) Slant range between the aircraft and the microphone:
SR(p) = [(X(p) – XMIC)2 + (Y(p) – YMIC)2 + (Z(p) – (ZMIC + HMIC))2]0.5
2) Sound propagation time:
δtprop(p) = SR(p) / c
3) Reception time:
trec(p) = t(p) + δtprop(p)
c) The emission/reception array should now include the following for each aircraft position:
t(p), X(p), Y(p), Z(p) and trec(p) Using linear interpolation, determine the time of sound emission, t E(k), for each kth measured acoustic spectral data record in the spectral time history:
tE(k) = t(p2) + ([ t(p1) - t(p2)] ([tm(k) - trec(p2)] / [trec(p1) - trec(p2)])) where:
— tE(k) is the relative time when spectrum k was emitted;
— tm(k) is the acoustic measurement time for spectral record k (as per Section 3.7.6 of Appendix 2 of the Annex) — p1 is the aircraft position record where trec(p) is > tm(k); and — p2 is the aircraft position record where trec(p) is < tm(k).
Once the time of sound emission has been determined generate the sound emission coordinates XE(k), YE(k) and ZE(k) for each kth measured acoustic spectral data record as follows:
XE(k) = (tE(k) – tOH) (VG cos(χ)) + XOH YE(k) = (tE(k) – tOH) (VG sin(χ)) + YOH ZE(k) = (tE(k) – tOH) (VG tan(γ))+ ZOH 403(a)(2)(E) Calculation of the geometric mi nimum distance between the test flight path and the microphone [ETM 4.3.1.2.5] Using the single-point flight path descriptors identified in 403(a)(2)(C) and the coordinates of the microphone identified in 403(a)(2)(D), calculate the geometrical minimum distance (closest point of approach or “CPA”), the line from the microphone of interest which intersects the straight line flight path at right angles:
CPA = (Gnorm2 + GCPA2)0.5 Intermediate calculations for determination of CPA (see Figures 4-31 and 4-32) include:
a) Lateral distance from micropho ne to average ground track at tOH:
Ydis = YMIC – YOH
b) Line on the ground from the microphone location which intersects the average ground track at right angles:
Gnorm = (Ydis) cos (χ
c) Vertical distance of flig ht path above microphone:
Malt = ZOH – (ZMIC + HMIC)
d) Distance along average ground track between X=0.0 and intersection with Gnorm:
Ginc = (Ydis) sin (χ
e) Vertical difference between G alt and ZOH:
Zinc = (Ginc) tan (γ
f) Vertical height of flight path above microphone at intersection of Gnorm and average ground track Galt = Malt + Zinc Page 197
g) Line from the point at the microphone height (H MIC) vertically above the point where Gnorm intersects the average ground track, which intersects the average straight line flight path at right angles GCPA = (Galt) cos (γ Figure 39. Basic CPA geometry Note.— In contrast to the previous figure, the horizontal X-Y plane illustrated in this figure is now at the height of the microphone, not the reference ground plane where Z=0.0.
Figure 40. Detailed CPA geometry Note.— This figure illustrates the CPA geometry for a lateral microphone, but also includes intermediate elements used in the calculation of Gnorm and G CPA applicable to a centerline microphone. Also, the X-Y plane in this figure is at the height of the lateral microphone.
Figure 41. Sound propagation distance, SR( k) Page 199
403(a)(2)(F) Determi nation of test aircraft noise geometry (sound propagation distance and acoustic emission angle for each acoustic s p e c t r u m ) [ETM 4.3.1.2.6] Using the sound emission coordinates from 403(a)(2)(D) calculate slant range SR(k), the sound propagation distance between the aircraft and the microphone, for each spectrum k in the spectral time history (see Figure 41):
SR(k) = [(XE(k) – XMIC)2 + (YE(k) – YMIC)2 + (ZE(k) – [ ZMIC + HMIC])2]0.5 Using the sound propagation distance, SR(k), and the geometric minimum distance between the flight path and microphone, CPA, from 403(a)(2)(E), calculate the three-dimensional acoustic emission angle, θ, for each spectrum k (see Figure 42):
θ(k) = arcsin (CPA / SR(k)) when aircraft is positioned prior to CPA θ(k) = π/2 (i.e. 90 degrees) when aircraft is positioned at CPA θ(k) = π – arcsin (CPA / SR(k)) when aircraft is positioned subsequent to CPA The resulting noise geometry time history for each acoustic spectral data record k now includes:
— tm(k) is the acoustic measurement time for spectrum k;
— tE(k) is the sound emission time for spectrum k, received at the microphone at tm(k);
— XE(k) is the X-coordinate at time of sound emission for spectrum k;
— YE(k) is the Y-coordinate at time of sound emission for spectrum k;
— ZE(k) is the Z-coordinate at time of sound emission for spectrum k;
— SR(k) is the sound propagation distance (slant range) for spectrum k; and — θ(k) is the three-dimensional acoustic emission angle for spectrum k.
Note.— SR(k) is used for calculations of sound attenuation due to spherical spreading, as well as for sound attenuation due to atmospheric absorption, and for the distance-dependent portion of the Δ2 duration adjustment term used in the simplified method; θ(k) is used for determining the aircraft position on the reference flight path when adjusting noise data to reference conditions.