4) quality of data to be extrapolated;
c) airplane engine and nacelle configuration and acoustical treatment changes, usually leading to changes in the values of EPNLr of less than 1 dB;
Note 1.— It should however be ensured that new noise sources are not introduced by modifications made to the airplane, engine or nacelles. A validated analytical noise model approved by the FAA may be used to derive predictions of noise increments. The analysis may consist of modelling each airplane component noise source and projecting the sources to flight conditions in a manner similar to the static test procedure described in 402(a)(3)(D). A model of detailed spectral and directivity characteristics for each airplane noise component may be developed by theoretical and/or empirical anal ysis. Each component shoul d be correlated to the parameters which relate to the physical behavior of source mechanisms. The source mechanisms, and subsequently the correlating parameters, should be identified through use of other supplemental tests such as engine or component tests. As described in 402(a)(3)(D), an EPNLr value representative of flight conditions should be computed by adjusting airplane component noise sources for forward speed effects and for the number of engines and shielding, reconstructing the total noise spectra, and projecting the total noise spectra to flight conditions by accounting for propagation effects. The effect of changes in acoustic treatment, such as nacelle lining, may be modeled and applied to the appropriate component noise sources. The computation of the total noise increments, the development of the changed version NPD, and the evaluation of the changed version EPNLr values should be made by using the procedures described in 402(a)(3)(D). Guidance material on confidence interval computations is provided in 305.
Note 2.— Some engine or aircraft design changes may change the nature of the noise signature of the engine or aircraft and prevent an ordered extension of noise data. In this case, static engine noise tests may not provide adequate proof that an extension of the noise database is valid. Changes that may cause transition of the fan tip velocity to supersonic, interaction of the mid-span fan shrouds or stator vanes, choking of the fan exit guide vanes or primary compressor entrance, operation of the surge bleed valve, an increase in the inlet bypass airflow or a change in aircraft configuration interaction are examples of such changes.
d) airframe design changes (e.g., changes in fuselage le ngth, flap configuration and engine installation) that could indirectly affect noise levels because of an effect on airplane performance (e.g., increased drag).
Note 1.— Changes in airplane performance characteristics derived from aerodynamic analysis or testing have been used to demonstrate how these ch anges affect the airplane flight path and hence the demonstrated noise levels of the airplane.
Note 2.— In these cases care should be exercised to ensure that the airframe design changes do not introduce significant new noise sources or modify existing source generation or radiation characteristics. In such instances the magnitude of such effects may have to be established by test.
402(a)(2)(C) Equivalent procedure for ca lculating the certification noise levels of weight variants of a given airplane type [ETM 4.2.1.2.3] Section 901 of Chapter 9 of this AC speci fies that “In the case of an airplane be ing recertificated to the standard of part 36, Amendment 24, noise certification should be granted on the basis that the evidence used to determine compliance is as satisfactory as the evidence expected of a new type design.” The lateral, flyover and approach noise levels and their 90 percent confidence intervals for the weight variants of a given airplane /engine model and acoustic configuration are typically derived from generalized NPD curves, based on the information in noise certification reports and supporting documentation, and in conjunction w ith certificated aerodynamic performance data for the airplane, as approved by FAA.
Some airplane manufacturers have used the noise level information initially certificated for several weight variants to demonstrate that when the basic airplane performance parameters (e.g., V2 and VREF ) vary in a linear manner over a range of certificated takeoff or approach weight, the resulting noise (EPNL) versus weight relationship can be shown Page 169
to be linear in that range as well. When this situation is demonstrated by the applicant and, subject to the approval by FAA, the applicant may derive the certification noise levels of additional weight variants using linear interpolation between previously certificated points. The confidence interval for the interpolated weight is then to be established in a process that utilizes the polynomial regression models that had been used by the applicant to develop the NPD. For a given airplane type, equivalency of the interpolated noise level values is demonstrated when the noise levels and the associated confidence intervals are calculated and reported in a manner acceptable to the FAA.
402(a)(3) Static engine noise tests and projections to flight noise levels [ETM 4.2.1.3] 402(a)(3)(A) General [ETM 4.2.1.3.1] Static engine noise test data provide valuable definitive information for deriving the noise levels that result from changes to an airplane powerplant or from the installation of a broadly similar powerplant into the airframe following initial noise certification of the “flight datum” airplane. This involves the testing of both the “flight datum” and derivative powerplants using an open-air test facility where the effect on the noise spectra of the engine modifications on airplane noise characteristics may be assessed. It can also extend to the use of component test data to demonstrate that when minor development changes have been made the noise levels remain unchanged (see section 101(d) of this AC).
Approval of equivalent procedures for the use of static engine noise test data depends critically upon the availability of an adequate approved database (NPD plot) acquired from the flight testing of the “flight datum” airplane.
Static engine noise tests can provide sufficient additional data or source noise characteristics to allow for predictions about the effect of changes on the airplane certification noise levels.
Types of static tests accepted for the purposes of demons trating certification compliance in airplane development include engine noise tests. Such tests are useful for assessing the effects on the individual noise sources of mechanical and thermodynamic cycle changes to the engine. Such configuration and/or design changes often occur as engines are developed, subsequent to the initial nois e certification of an aircraft, to eas e production difficulties, reduce cost, improve durability, and for operational reasons.
Static engine noise testing is discussed in detail in subsequent sections. For component tests, the criteria for acceptability are less definable. There are many instan ces, particularly when only small changes in EPNL r are expected, where component testing will provide an adequate demonstration of noise impact. Examples of such changes include:
a) changes in the specification of sound-abs orbing linings within an engine nacelle;
b) changes in the mechanical or aerodynamic design of the fan, compressor or turbine;
c) changes to combustor designs, including material changes;
d) changes to bleed valves; and
e) changes to the exhaust system.
Each proposal by an applicant to use component test data should be considered by the FAA with respect to the significance of the relevant affected source on the values of EPNLr for the airplane that is being certificated.
402(a)(3)(B) Limitation on the projection of static-to-flight data [ETM 4.2.1.3.2] Page 170
Guidance on the acceptability, use and applicability of static engine test data are contained in subsequent sections.
The amount by which the measured noise levels of a derivative engine will differ from those of the reference engine is a function of several factors, including:
a) thermodynamic changes to th e engine cycle, including increases in thrust (power);
b) design changes to major components (e.g., the fan, compressor, turbine, exhaust system); and
c) changes to the nacelle.
Additionally, day-to-day and test-site to test-site variables can influence measured noise levels, and therefore the test, measurement and analysis procedures described in this AC are designed to account for these effects. A limit is needed that can be used uniformly by FAA in order that the degree of change resulting from aspects such as a), b) and c), when extrapolated to flight conditions, is restricted to acceptable amounts before a new flight test is required.
The recommended guideline for this limit is that the summation of the magnitudes, neglecting signs, of the noise changes for the three reference certification conditions between the “flight datum” airplane and the derived version at the same thrust (power) and distance for the derived version is no greater than 5 EPNdB, with a maximum of 3 EPNdB at any one of the reference conditions (s ee Figure 31). For differences greater than this, additional flight testing at conditions where noise levels are expected to change is recommended in order to establish a new flight NPD database.
Provided that the detailed prediction procedures used are verified by flight tests for all the types of noise sources (i.e., tones, non-jet broadband and jet noise relevant to the airpla ne under consideration) and th at there are no significant changes in installation effects between the airplane used for the verification of the prediction procedures and the airplane under consideration, the procedure may be employed without the limitations described above.
In addition to the limitations described above, a measure of acceptability regarding methodologies for static-to-flight projection is also needed for uniform application by FAA. This measure can be derived as the residual NPD differences between the flight test data and the projected static-to-flight data for the original airplane version. The guideline for a measure of acceptability is to limit these residual differences to 3 EPNdB at any one of the reference conditions.
In determining the noise levels of the modified or derived version, the same analytical procedures used in the first static-to-flight calculations for the noise certification of the airplane type must be used.
Derivative aeroplane L, T or A Derivative aeroplane µ requirement Derived by extrapolation of datum flight-projected NPD where necessary Flight-projected EPNL from static test Engine noise performance parameter Definitions L = Difference between flight datum and derivative aeroplane EPNL at the power requirement of the derivative aeroplane at the lateral condition.
T = Difference between flight datum and derivative aeroplane EPNL at the power requirement and altitude of the derivative aeroplane at the take-off condition.
A = Difference between flight datum and derivative aeroplane EPNL at the power requirement of the derivative aeroplane at the approach condition.
Limitations
(i) |L| + | T| + | A| shall not exceed 5 EPNdB.
i) L T or A shall not exceed ± 3 EPNdB individually. (i , Figure 31. Limitation on use of static test when no validating flight data exist 402(a)(3)(C) Static engi ne noise test procedures [ETM 4.2.1.3.3] 402(a)(3)(C)(i) General [ETM 4.2.1.3.3.1] Test restrictions defined for flight testing in conformity with part 36 are not necessarily appropriate for static testing.
Reference T13 provides appropriate, detailed guidance for measurement of far-field sound pressure levels during static operation of gas turbine engines installed on an outdoor test stand. The following sections provide guidelines supplemental to sections on test site characteristics, data acquisition and reduction, microphone types, locations and installations provided in Reference T13. These supplemental guidelines are specifically related to static engine noise testing for purposes of airplane noise certification.
Noise data acquired from static tests of engines with similar designs to those that were flight tested may be projected to flight conditions when appropriate. Once approved, noise data acquired from static tests may be used to supplement an approved NPD plot for the purpose of demonstrating compliance with the part 36 provisions in support of a change in type design. The engine designs, as well as the test and analysis techniques to be used, should be presented in the test plan and submitted to the FAA for approval prior to testing.
402(a)(3)(C)(ii) Test site characteristics [ETM 4.2.1.3.3.2]
(1) Inflow control devices (ICD) Static engine noise test data for the noise certification of an airplane with a change of engine to another one of a Page 172
similar design should be acquired by using an approved ICD for high bypass engines (i.e., BPR > 2.0). The ICD should meet the requirements in Reference T13. The speci fic ICD hardware should be inspected by the FAA to ensure that the ICD is free from damage and contaminants that may affect its acoustic performance.
(2) ICD calibration An acceptable ICD calibration method is provided in Reference T13. In some cases large fluctuations in the value of the calibrations across adjacent one -third octave bands and between cl osely spaced angular positions of microphones can occur. These fluctuations can be related to reflection effects caused by the calibration procedure, and care must be taken to ensure that they do not introduce or suppress engine tones. This may be done by comparing EPNL computed with:
a) the ICD calibrations as measured;
b) a mean value of the calibration curves; and
c) the calibration values set to zero.
402(a)(3)(C)(iii) Data acquisition and reduction systems [ETM 4.2.1.3.3.3]
(1) Data acquisition, analysis and normalization Data acquisition and analysis systems used for static test and the modus operandi of the test program may well vary according to specific test objectives. In general they should conform to those outlined in Reference T13.
For each engine power setting designated in the test plan , the engine performance, meteorological and sound pressure level data should be acquired and analyzed using measurement systems and test procedures described in Reference T13 or as approved by the FAA.
A range of static engine operating conditions should be selected to correspond to the expected maximum range of in-flight engine operating conditions for the appropriate engine power setting parameter. A sufficient number of stabilized engine power settings over the desired range should be included in the test to ensure that the 90 percent confidence intervals for values of flight-projected EPNL can be established (see 305(c)).
Noise measurements should be normalized to consistent conditions and include 24 one -third octave band sound pressure levels between band center frequencies of 50 Hz to 10 kHz for each measurement microphone. Before projecting the static engine data to flight conditions, the sound pressure level data should be adjusted for:
a) the frequency response characteristic s of the noise measurement system; and
b) contamination by background or electrical system noise (see 306(c) of this AC).
(2) Data system compatibility If more than one data acquisition system and/or data analysis system is used for the acquisition or analysis of static data, compatibility of the airfra me and engine manufacturers’ system s is necessary. Compatibility of the data acquisition systems can be accomplished through appropriate calibration. Compatibility of the data analysis systems can be verified by analyzing the same flight test noise data samples on both systems. The systems can be considered to be compatible if the resulting differences are no greater than 0.5 EPNdB. Evaluation should be conducted at flight conditions representative of those for certification.
The use of measured static engine test noise data or pseudo-random noise signals with spectral shape and tonal content representative of turbofan engines is an acceptable alternative to the use of actual flight test noise data Page 173
samples for determination of analysis system compatibility. The systems can be considered to be compatible if the resulting differences are no greater than 0.5 PNdB for an integration time of 32 seconds.
402(a)(3)(C)(iv) Microphone types, locations and installations [ETM 4.2.1.3.3.4]
(1) Microphone locations Microphones should be located over an angular range sufficient to include the 10 dB-down times after projection of the static noise data to flight conditions. The general guidance in Reference T13 describing microphone locations is sufficient to ensure adequate definition of the engine noise source characteristics.
The choice of microphone location with respect to the test surface depends on the specific test objectives and the methods to be used for data normalization. Certification experience with static engine testing has been primarily limited to microphone installations near the ground or at engine centerline height. In general, because of the difficulties associated with obtaining free-field sound pressure levels that are often desirable for extrapolating to flight conditions, near-ground-plane microphone installati ons or a combination of ground-plane and elevated microphones have been used. Consistent microphone locations, heights, etc., are recommended for noise measurements of both the prior approved and changed version of an engine or nacelle.
(2) Microphone installations Where ground-plane microphones are used, special precautio ns as outlined in Reference T13 are necessary to ensure that consistent measurements (e.g., free from “ac oustic shadowing” refraction effects) will be obtained.
When there is a wind in the opposite direction to the sound wave propagating from the engine, or when there is a substantial thermal gradient in the test arena, refraction can influence near-ground-plane microphone measurements to a larger degree than measurements at greater heights.
402(a)(3)(D) Projection of static engine data to airplane flight conditions [ETM 4.2.1.3.4] 402(a)(3)(D)(i) General [ETM 4.2.1.3.4.1] The static engine sound pressure level data acquired at eac h angular location should be analyzed and normalized to account for the effects identified in the pa ragraphs below. They should then be projected to the same airplane flight conditions used in the development of the approved NPD plot.
As appropriate, the projection procedure includes the:
a) effects of source motion including Doppler effects;
b) number of engines and shielding effects;
c) installation effects;
d) flight geometry;
e) atmospheric propagation, including spherical wave divergence and sound attenuation; and
f) flight propagation effects, including ground reflection and lateral attenuation.
To account for these effects, the meas ured total static noise data should be analyzed to determine contributions from individual noise sources. After projection of the one-third octave band spectral data to flight conditions, Page 174
EPNLs should be calculated for the revised NPD plot. Guid elines on the elements of an acceptable projection procedure are provided in this section. The process is also illustrated in Figures 4-24 and 4-25.
It is not intended that the procedure illustrated in Figu res 4-24 and 4-25 should be exclusive. There are several options, depending upon the nature of the powerplant noise sources and the relevance of individual noise sources to the EPNL of the airplane. The method presented does however specify the main features that should be considered in the computational procedure.
It is also not necessary that the computations illustrated in Figures 4-24 and 4-25 should always be carried out in the order specified. There are interrelations between th e various steps in the procedure which depend on the particular form of the computation being followed. Hence the most efficient manner of structuring the computation cannot always be predetermined.
Combine to give total aeroplane noise and adjust for propagation effects Compute EPNL Extrapolate to aeroplane flight path Account for airframe noise Adjust for aeroplane configuration Adjust for noise source motion and engine flight conditions Adjust for noise source position Separate into contributing noise sources Separate into broadband and tone noise Correct for test environment Acquire static noise data Broadband Tones Normalized flight spectra Flight noise source breakdown Static noise source breakdown Normalized freefield static spectra Figure 32. Generalized projection of static engine data to airplane flight conditions (Refer to 402(a)(3)) Page 175
Acquire time-averaged static test data under stabilized power conditions over angular range embracing 10 dB-down times of flight time-history.
For each individual measured angle* perform one-third octave (and possibly narrow-band) analysis and adjust data for static propagation effects under reference atmosphere conditions.
Identify all relevant discrete tones and subtract from one-third octave spectrum to give broadband spectrum.
Consider tones and broadband separately.
Broadband Tones Using approved source identification techniques, conduct source breakdown.
Adjust for source motion effects, accounting for relative velocity, Doppler frequency shift, convective amplification, directivity effects, etc., and engine thermodynamic changes.
Adjust for number of engines, relevant installation effects and sources of interaction/airframe noise, and reconstitute one-third octave band sound pressure levels.
Extrapolate to aircraft height (slant distance) and allow for ground reflection, lateral propagation and other relevant effects to produce quasi in-flight one-third octave band SPL and PNLT.
Integrate projected PNLT over time corresponding to height, flight path and airspeed.
Compute EPNL Identify origin of tones where appropriate.
PNLT* Narrowband SPL SPL One-third octave Frequency Frequency Jet Other Fan Aeroplane Fan tones Airframe PNLT 10 dB Time Repeat for other angles EPNL Figure 33. Example procedure for projection of static engine data to airplane flight conditions There are several engine installation effects which can modify the generated noise levels but which cannot be derived from static tests. Additional noise sources such as jet/flap or jet/wind interaction effects may be introduced on a derived version of the airplane which are not present on the “flight datum” airplane Far-field noi se directivity patterns (i.e., field shapes) may be modified by wing/nacelle or jet-by-jet shielding, tail plane and fuselage scattering or airframe reflection effects. However general methods to adjust for thes e effects are not yet available. It is therefore important that before the following procedures are approved for the derived version of the airplane, the geometry of the airframe and engines in the vicinity of the engines be shown to be essentially identical to that of the “flight datum” airplanes so that the radiated noise is essentially unaffected.
402(a)(3)(D)(ii) Normalizat ion to reference conditions [ETM 4.2.1.3.4.2] Page 176
The analyzed one-third octave band sound pressure level static test data should be normalized to free-field reference atmospheric conditions specified in B36.7(a)(k)(5) of part 36.This adjustment can be applied only with knowledge of the total spectra being the summation of all the noise so urce spectra computed as described in the following paragraphs. The required adjustments include:
a) Sound attenuation due to atmospheric absorption . Adjustments to account for the acoustical reference-day sound attenuation are defined in Reference T16. In the event that minor differences in coefficient values are found in Reference T16 between equations, tables or graphs, the equations should be used. The sound attenuation coefficients should be computed over the actual distance from the effective center of each noise source to each microphone, as described in 402(a)(3)(D)(v) of this AC.
b) Ground reflection. Examples of methods for obtaining free-field sound pressure levels are described in References T5 and T15. Spatial distribution of noise sources do not have a first order influence on ground reflection effects and hence may be disregarded. It is also noted that measurements of far-field sound pressure levels with ground-plane microphones may be used to avoid the large spectral irregularities caused by interference effects at frequencies less than 1 kHz.
402(a)(3)(D)(iii) Separation into broadband and tone noise [ETM 4.2.1.3.4.3] The purpose of procedures described in this section is to id entify all significant tones in the spectra: first, to ensure that tones are not included in the subsequent estimation of broadband noise and, second, to enable the Doppler-shifted tones in-flight to be allocated to the correct one-third octave band at appropriate times during a simulated airplane flyover.
Broadband noise should be derived by extracting all significant tones from the measured spectra. One concept for the identification of discrete tones is the one used in Appendix A of part 36 for tone correction purposes (i.e., considering the slopes between adjacent on e-third octave band levels). Care must be taken to avoid regarding tones as “non - protrusive” when the surrounding broadband sound pressure level is likely to be lower when adjusted from static-toflight conditions, or when classifying a closely grouped pair or series of tones as broadband noise. One technique for resolving such problems is the use of narrow-band analysis with a bandwidth of less than 50 Hz.
Narrow-band analysis can also be used to check the validity of other tone identification procedures in establishing the spectral character at critical locations in the sound field (e.g., around the position of peak PNLT) or where predominant turbo-machinery tones exist.
402(a)(3)(D)(iv) Separation in to contributing noise sources [ETM 4.2.1.3.4.4] The number of noise sources which require identification will to some extent depend on the engine being tested and the nature of the change to the engine or nacelle. The separation of broadband noise into the combination of noise generated by external jet mixing and by internal noise sources is the minimum and sometimes adequate requirement. A more sophisticated analysis may be necessary depending upon th e significance of the contri bution from other individual sources, which could involve identifying broadband noise from fan, compressor, combustor and turbine. Furthermore, for fan and compressor noise, the split of both the broadband and the tone noise between that radiating from the engine intake and that from the engine exhaust nozzle(s) could be a further refinement.
To meet the minimum requirement, the separation of sources of broadband noise into those due to external jet mixing and those generated internally can be carried out by:
a) estimating the jet noise by one or more of the methods identified below; and
b) adjusting the level of the predicted spectrum at each angle to fit the measured low frequency part of the broadband spectrum at which jet noise can be expected to be dominant.
There are three methods which have been used to obtain predicted jet noise spectra shapes:
a) for single-stream engines with circular nozzles, th e procedure detailed in Reference T12 may be used; the engine geometry however may possess features which can render this method inapplicable; sample procedures for coaxial flow engines are provided in Reference T11;
b) analytical procedures based on correlating full-scale engine data with model nozzle characteristics may be used; model data have been used to supplement full-scale engine data, particularly at low power settings, because of the uncertainty in defining the level of jet noise at the higher frequencies where noise from other engine sources may make a significant contribution to the broadband noise; and
c) special noise source location techniques are available which, when used during full-scale engine tests, can identify the positions and levels of separate engine noise sources.
402(a)(3)(D)(v) Noise source position effects [ETM 4.2.1.3.4.5] Static engine noise measurements are often made at distances at which engine noise sources cannot be truly treated as radiating from a single acoustic center. This may not give rise to difficulties in the extrapolation to determine the noise increments from static data to flight conditions because noise increments in EPNL are not particularly sensitive to the assumption made regarding the spatial distribution of noise sources.
However, in some circumstances, for example where changes are made to exhaust structures and where the sources of external jet-mixing noise are of overriding significance, it may be appropriate to identify noise source positions more accurately. The jet noise can be considered as a noise source distributed downstream of the engine exhaust plane.
Internal sources of broadband engine noise may be considered as radiating from the intake and the exhaust.
There are three principal effects to be accounted for as a consequence of the position of the noise source differing from the “nominal” position assumed for the “source” of engine noise:
a) Spherical divergence. The distance of the source from the microphone differs from the nominal distance, in which case an inverse square law adjustment needs to be applied.
b) Directivity. The angle subtended by the line from the source to the microphone and the source to the engine centerline differs from the nominal angle, in which case a linear interpolation should be made to obtain data for the proper angle.
c) Sound attenuation due to atmospheric absorption. The difference between the true and the nominal distance between the source and the microphone alters the allowance made for sound attenuation.
Source position can be identified either from noise source location measurements (made either at full or model scale) or from a generalized database.
Note.— No published standard on coaxial jet noise source distribution is currently available. An approximate distribution for a single jet is given by the following equation (see References T6 and T17):
x/D = (0.0575 S + 0.0215 S–2 )–½ where:
— S is the Strouhal number fD/V i;
— x is the distance downstream from the nozzle exit;
— D is the nozzle diameter based on total nozzle exit area;
— Vj is the average jet velocity for co mplete isentropic expansion to ambient pressure from average nozzleexit pressure and temperature; and — f is the one-third octave band center frequency.
402(a)(3)(D)(vi) Engine flight conditions [ETM 4.2.1.3.4.6] Some thermodynamic conditions within an engine tested sta tically differ from those that exist in flight and this difference should be taken into account. Noise source strengths may be changed accordingly. Therefore, the values for key correlating parameters for component noise source generation should be based on the flight condition, and the static database should be entered at the appropriate correlating parameter value. Turbo-machinery noise levels should ேభbe based on the in-flight corrected rotor speeds √ఏమ Jet noise levels should be based on the relative jet velocities that exist at the flight condition.
The variation of source noise levels with key correlating parameters can be determined from the static database which includes a number of different thermodynamic operating conditions.
402(a)(3)(d)(vii) Noise source motion effects [ETM 4.2.1.3.4.7] The effects of motion on jet noise differ from speed effects on other noise sources and, hence, are considered separately during static-to-flight projection.
(a) For external jet noise Account should be taken of the frequency-dependent jet-relative velocity effects and the convective amplification effects. Broadly speaking, two sources of information may be used to develop an approved method for defining the effect of flight on external jet noise:
1) for single-stream engines with circular exhaust geometries, Reference T12 provides guidance. Additional supporting evidence however may be needed to show when jet noise is the major contributor to the noise from an engine with a more complex nozzle assembly; and
2) full-scale flight data on a similar exhaust geomet ry can provide additional evidence. In general however, because of the difficulty of defining high frequency effects in the presence of internally-generated engine noise, it may be necessary to provide additional sup porting information to determine the variation of EPNL with changes of jet noise spectra at high frequencies.
(b) For noise sources other than jet noise In addition to the Doppler frequency effect on the non-jet noise observed on the ground from an airplane flyover, the noise generated by the engine’s internal components and the airframe can be influenced by source amplitude modification and directivity changes:
1) Doppler effect. Frequency shifting that results from motion of the source (i.e., airplane) relative to a microphone is accounted for by the following equation:
݂௦௧௧݂௧ ൌ ሺ1 െ ߣݏܿܯ Page 179
w h e r e :
— ݂௧ is the flight frequency;
— ݂௦௧௧ is the static frequency;
— M is the Mach number of the airplane; and — is the the angle between the flight path in the direction of flight and a straight line connecting the airplane and the microphone at the time of sound emission.
It should be noted for those one-third octave band sound pressure levels dominated by a turbo-machinery tone, the Doppler shift may move the tone and its harmonics into an adjacent band.
2) Source amplitude modification and directivity changes. One-third octave band sound pressure level adjustments to airframe-generated noise that re sults from speed change s between the datum and derivative versions provided below.
For noise generated internally within the engine (e.g., fan noise), there is no consensus of opinion on the mechanisms involved or on a unique adjustment method that accounts for the detailed source modification and sound propagation effects. If an adjustment is used, the same technique must be applied to both the flight datum and derivative configuration when establishing noise changes. In such instances the adjustment for the one-third octave band sound pressure level changes that result from the motion of the source (i.e., airplane) relative to the microphone may be accounted for by using the following equation:
ܮܲܵ௧ ൌܮܲܵ ௦௧௧ െ݈݃ܭ ሺ 1െ ߣݏܿܯ ሻ where:
— SPL flight is the flight sound pressure level;
— SPL static is the static sound pressure level; and — M and are defined above and K is a constant.
Theoretically, K has a value of 40 for a point noise source, but a more appropriate value may be obtained by comparing static and flight data for the “flight datum” airplane.
402(a)(3)(D)(viii) Airplane configuration effects [ETM 4.2.1.3.4.8] The contribution from more than one engine on an airplane is normally taken into account by adding 10 log N, where N is the number of engines, to each component noise source. It might be necessary however to compute the noise from engines widely spaced on large airplanes, particularly in the approach case, if they include both underwing and fuselage mountings. The noise from the intakes of engines mounted above the fuselage is known to be shielded.
If engine installation effects change between the “flight datum” airplane and a derived version, account should be taken of the change on one-third octave band sound pressu re levels which should be es timated according to the best available evidence.