AC 36-4B 3/23/88 FIGURE 3
POSSIBLE APPLICATIONS OF SUPPLEMENTAL
NOISE DEMONSTRATION METHODS
EXAMPLE OF CHANGES IN TYPE DESIGN
AIRPLANE AERO-PERFORMANCE CHANGES SISA TUNG (e.g., Gross Weight, Flap, Fuselage, etc.)
MINOR ENGINE THRUST AND/OR RATING
CHANGES
MINOR MODIFICATIONS TO ENGINE
MINOR MODIFICATIONS TO ENGINE
NACELLE
COMPONENT TESTING ENGINE COMPONENTS (e.g., Jet, Fan, etc.)
BOUNCE •
ACOUSTIC TREATMENT CHANGES
NEW ENGINE OF COMPARABLE TECHNOLOGY
AND SOURCE NOISE CHARACTERISTICS
TESTING STATIC ACOUSTIC TREATMENT CHANGES
NEW ENGINE NACELLE
ENGINE COMPONENT CHANGES SUPPLEMENTAL FLIGHT NOISE SUPPRESSION DEVICE (e.g., TESTING Exhaust Gas Mixer) ENGINE COMPONENTS (e.g., Jet, Fan, etc.)
NOTE: THIS LIST IS NOT COMPREHENSIVE AND
FOR A SPECIFIC TYPE CHANGE, MORE
THAN ONE DEMONSTRATION METHOD MAY BE TECHNICALLY VALID.
Chap 3 23 Par 16
This avoids discrepancies that can occur from merging differing methods.
demonstration method is designed to provide noise levels representative of those occurring with production hardware (e.g., engine and nacelle).
Furthermore, the spectral characteristics of the basic and modified configurations need to be similar.
d. Supplemental Test Data Compatibility. The application of a noise demonstration method depends upon the compatibility with data previously acquired. Items which should be considered include test hardware, test and analysis procedures, test site and configuration.
e. Component Noise Sources.
affected by the change should be considered in determination of the best method for the evaluation of changes to the total airplane source noise characteristics. Parameters controlling the component noise source generation nechanisms and directional characteristics should be identified in order to assess the noise sensitivity of each component source. For example, when the total airplane noise characteristics are dominated by one source, such as the low speed rotor, the use of component test results, in combination with a component noise source analysis, may be an appropriate method to demonstrate
f. Noise Increments Caused by Nacelle Acoustic Treatment Changes. Noise increments caused by nacelle acoustic treatment changes made to noise certified aircraft can be empirically derived by coupling data from static engine noise tests with data from certification flyover noise tests. The noise increments, in EPNdB, can be determined by applying the difference in measured spectra between the two nacelle treatments to measured one-third octave band sound pressure levels in flight at the same directivity angle.
the modified nacelle treatment produces levels lower than the existing nacelle treatment, noise generated by the airframe should be considered in determining the dominant noise source.
17. STATIC ENGINE TESTS. Data acquired from static tests of engines incorporating similar designs to those that were flight tested may be projected, where appropriate, to flight conditions and, after FAA approval, used to supplement an approved NPD for the purpose of demonstrating compliance with Part 36 in support of a change in type design. This section provides guidelines on static engine test data acquisition, analysis and normalization techniques. The information provided is used in conjunction with the technical considerations in paragraph 16 and the general guidelines for test site, measurement and analysis instrumentation and test procedures provided in Society of Automotive Engineers (SAE) Aerospace Recommended Practice (ARP) No. 1846, Measurement of Noise from Gas Turbine Engines During Static Operation. The engine designs and the test and analysis techniques to be used should be presented in the test plan and submitted to the FAA for concurrence prior to conducting the test. It should be noted that test restrictions defined for flight testing in conformity with Part 36 are not necessarily appropriate for static testing. For example, the distances associated with atmospheric absorption for static tests are substantially less than those encountered in flight testing and thus may require different criteria.
Chap 3 24 Par 16
Limits on the Projection of Static to Flight Data. The amount by which the measured noise levels of a derivative engine will differ from the reference engine is a function of several factors, including: (1) thermodynamic changes to the engine cycle, including increases in thrust; (2) design changes to major components, e.g., the fan, compressor, turbine, exhaust system, etc.; and (3) changes to the nacelle. Additionally, day-to-day and test site-to-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. While the FAA does not impose a fixed numerical limit on the amount of extrapolation from reference to derivative engines, it 1s important to consider whether there are practical limitations to proposed extrapolation techniques and at what point new flight test data may be required.
Test Site Requirements. The test site should meet at least the criteria specified in SAE ARP 1846. Different test sites may be selected for testing the subject configurations provided the test sites approved by the FAA are acoustically similar. See paragraph 17(g) of this AC for criteria concerning data acquisition and analysis systems. Depending upon test objectives and unique test stand and test site characteristics, wind speed and direction limits should be considered.
c. Test Hardware. A flight inlet adjusted for static testing by the addition of a bellmouth forward of the inlet may be used with each turbofan or turbojet static noise test. Production inlet acoustic lining and spinners are also to be installed during noise testing.
d. Inflow Control Devices. The use of static engine test noise data for Part 36 noise certification under an approved "family plan" requires the use of an approved Inflow Control Device (ICD) for high bypass engines (BPR 2.0).
For static tests conducted after May 1, 1986, the following requirements have been imposed on the ICD:
(1) The specific ICD hardware must be inspected for conformance with an approved ICD design by an approved method.
(2) The ICD must be acoustically calibrated for sound transmission loss by an approved method.
(3) The ICD acoustic calibration must be used as a correction in the certification data analysis. In applying the calibration, an approved method must be used to determine that portion of the noise received at the forward quadrant microphones that 1s radiated through the ICD. The calibration is applied to that portion of the noise that is radiated through the ICD.
(4) The ICD position relative to the engine inlet must be determined and found to be one of the calibrated positions.
(5) The ICD calibration consists of an SPL correction for each one-third octave band at each forward quadrant microphone that will be used during static engine testing.
Chap 3 Par 17 25
(6) No more than one calibration is required for an ICD hardware design, provided that there is no deviation from the design for any one ICD serial number hardware set.
The above calibration requirements are not required if the same ICD hardware (identical serial number) is used as was previously used in the static noise test of the flight engine configuration, and the fan tones for both engines remain in the same one-third octave bands.
e. ICD Calibration. An acceptable ICD calibration method is as follows:
(1) Assume the correction is zero for the one-third octave bands from 30 Hz through 800 Hz, unless the engine design would dictate that forward radiated tones or major inlet broadband noise occurs below 1000 Hz. In that case, an approved correction procedure is required.
(2) Place an acoustic driver(s) on a simulated engine centerline in the plane of the engine inlet l1p. Locate the calibration outer microphones on the forward quadrant azimuth at a radius between 50 and 150 feet that provides a good signal-to-ambient noise ratio and at each microphone angle to be used to analyze static engine noise data. Locate a near-field microphone on the centerline of and within 2 feet of the acoustic center of the acoustic driver (s).
(3) Conduct a test without the ICD in place. The test must be for a minimum of 60 seconds duration following system stabilization. The test must be conducted at a constant input voltage to the acoustic driver (s).
(4) Repeat item (3), alternately with and without the ICD in place.
A minimum of three tests of each configuration is required. To be acceptable, the total variation of the 55° microphone on-line OASPL signal (averaged for a l minute duration) for all three test conditions of each configuration shall not exceed 0.5 dB. NOTE: Physically moving the ICD alternately in and out of place for this calibration may be eliminated if it is demonstrated that the ICD positioning does not affect the calibration results.
(5) All measured data are to be corrected for sound pressure level variations as measured with the near-field microphone and for atmospheric absorption to 77°F and 70 percent RH conditions using the slant distance between the outer microphones and the acoustic driver (s).
(6) The calibration for each one-third octave band at each microphone is the difference between the average of the corrected SPLs without the ICD in place and the average of the corrected SPLs with the ICD in place.
(7) The tests must be conducted under wind and thermal conditions that preclude acoustic shadowing at the outer microphones and weather induced variations in the measured SPL data.
f. Acoustic Shadowing.
(1) If ground-plane microphones are used, the applicant should demonstrate that acoustic "shadowing" resulting from either thermal gradients Chap 3 26 Par 17
or wind does not significantly influence the measured sound pressure levels.
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 area, refraction can influence near-ground plane microphone measurements to a larger degree than measurements at greater heights. Previous evidence, or data from supplemental tests, may be used to demonstrate that testing within the approved weather restrictions at a particular test site results in no significant evidence of shadowing. In lieu of this evidence, the supplemental noise demonstration tests should include an approved method to indicate the absence of shadowing effects. One such method that has been approved involves the use of three measurements located as follows:
- average windspeed at engine centerline height (WL)
- air temperature at engine centerline height (TCL)
- air temperature at the ground-plane microphone height (TIC)
(2) Ground plane microphones should be mounted within one-half microphone diameter or one-quarter wavelength of the highest frequency to be recorded, whichever is least. Generally, this spacing is 7 mm (0.25 inch).
(3) The instruments for these measurements should be co-located at a distance from the engine similar to that of the microphones and placed close to the 90 degree noise measurement position without impeding the acoustic measurement. The limits that follow are in addition to any wind and temperature limits established by other criteria, such as the maximum crosswind at the engine and the maximum windspeed at the microphones if microphone windscreens are not used. For an acceptable data test point, the average windspeed and temperature should be within the following limits:
- Engine centerline height at 16 ft.
(assuming a 150 ft radius microphone array and smooth concrete pad):
(ICL - IMIC) ≥ -7°F (-3.9°C) , WCL≤11 km/h (6 kts) average during test time for each data point (typically 30 seconds), WCL≤15 km/h (8 kts) maximum during test time for each data point.
- Engine centerline height at 21 ft.
(assuming a 150 ft radius microphone array and smooth concrete pad):
(ICL - TMIC) = -7°F (-3.9°C), Chap 3 Par 17 27
WCLE 19 km/h (10 kts) average during test time for each data point, WCL≥ 23 km/h (12 kts) maximum during test time for each data point.
Microphone Locations. Microphones should be located at sufficient angular locations and at a height to provide adequate definition of the engine noise source characteristics. More specific guidance is given in SAE ARP 1846. The objective is to obtain measurements which can be corrected to a common reference condition. One method is to obtain or correct data to free field conditions (see paragraph 18(b) of this AC). Certification experience has been primarily limited to microphone installations near the ground or at centerline height. Because of the difficulties associated with extrapolating to flight conditions, the FAA suggests using near-ground plane microphone installations. However, the choice of microphone installations is dependent upon the specific test objectives and FAA approval of methods to be used for data normalization. It is currently required to use consistent microphone locations, heights, etc., for noise measurements of both the prior approved and changed aircraft version.
h. Engine Power Test Conditions. A range of static engine operating conditions should be selected to correspond to the expected maximum range of in-flight engine operating conditions. A sufficient number of engine power settings should be tested to ensure that the 90 percent confidence intervals in flight projected EPNL can be established over the desired range. (See Appendix 5 of this AC.)
1. Data Analysis System Compatibility. If more than one data analysis system is used for the acquisition of static data, compatibility of the two analysis systems is necessary. This can be verified by analyzing the same tape on both systems. The systems are compatible if the resulting differences in the projected in-flight EPNL values are no greater than 0.5 EPNdB.
Evaluation should be conducted at flight conditions representative of those for certification.
J. Data Acquisition, Analysis, and Normalization. For each power setting designated in the test plan, the engine performance, meteorological, and sound pressure level data should be acquired and analyzed. Instrumentation and test procedures described in SAE ARP 1846 are acceptable. Sound measurements should be normalized to consistent conditions (generally free field) and consist of 24 one-third octave band RMS sound pressure levels between 45 and 11,200 Hz for each measurement (microphone) station. Before projecting the static engine data to flight conditions, the sound pressure level data should be corrected for the following effects:
(1) Frequency response characteristics of the data acquisition and analysis system.
(2) Spherical divergence over the sound propagation path.
Chap 3 28 Par 17
(3) Atmospheric absorption over the sound propagation path due to the difference in test and Part 36 reference ambient conditions of temperature and relative humidity.
(4) Spectral distortions resulting from ground reflections (methods to account for spectral distortions are presented in SAE Aerospace Information Report (AIR) No. 1672B).
(5) Contamination by background ambient or electrical system noise.
18. PROJECTION OF STATIC ENGINE DATA TO AIRPLANE FLIGHT CONDITIONS. The static engine sound pressure level data acquired at each angular location should be analyzed and adjusted to account for the effects identified in paragraph 17(h) of this AC. They should be then projected to the same airplane flight conditions used in the development of the approved NPD. As appropriate, the projection procedure includes the effects of source motion, number of engines and shielding, flight geometry and atmospheric propagation.
To account for these effects, the measured static noise data should be analyzed to determine contributions from individual noise sources. After projecting the one-third octave band spectral data to flight conditions, EPNL values should be calculated for the revised NPD plot. Guidelines on the elements of an acceptable projection procedure are provided in this section.
This process is also illustrated in Figures 4 and 5. It is not intended that the procedure illustrated in Figures 4 and 5 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.
method presented does, however, specify the main features that should be considered in the computational procedure. It is also not necessary that the computations should always be carried out in the order specified. There are interrelations between the 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.
NOTE: Static-to-flight projection procedures for installations or engine types that are not similar to those that have been previously approved may require flight validation.
Engine Installation Effects. There are several engine installation effects which can modify the generated noise levels but which are not derivable from static tests. Additional noise sources, such as jet/flap or jet/wing interaction effects, may be introduced on a derivative version of the airplane which are not present on the flight datum airplane. The aircraft's noise directivity patterns (field shapes) may be modified by wing/nacelle or jet-by-jet shielding, tailplane and fuselage scattering or airframe reflection effects. However, general methods to adjust for these effects are not yet available. It is, therefore, important that, before the following procedures are approved for the derivative version of the airplane, the geometry of the airframe in the vicinity of the engines can be shown to be essentially identical to that of the flight datum airplane so that the radiated noise is unaffected.
Chap 3 Par 18 29
FIGURE 4
EXAMPLE OF A METHOD FOR PROJECTING STATIC ENGINE
DATA TO FLIGHT CONDITIONS
STATIC DATA T (EACH ANGLE) NOISE SECTRA
SEPARATE INTO CONTRIBUTING SOURCES
COMPONENT
OTHER SOURCE JET (FAN. COMPRESSOR SEPARATION COMBUSTOR. TURBINE ETC.)
ADJUST FOR SOURCE MOTION * JET REL. * DOPPLER VELOCIT\ EFFECTS
EFFECTS
PROJECT TO PROPER DISTANCE PROJESTION * SPHERICAL DIVERGENCE FLIGHT * ATMOSPHERIC ABSORPTION
ADJUST FOR FLIGHT GEOMETRY * NO. OF ENGINES * SHIELDING *1.2M MICROPHONE HEIGHT (APPROACH AIRFRAME NOISE COMBINE FOR TOTAL NOISE (LOG SUM COMPONENTS) TOTAL PROJECTED FLIGHT SFECTRA NOISE ONLY)
CALCULATE EPNL 30 Chap 3 Par 18
FIGURE 5
PROJECTION OF STATIC ENGINE DATA TO AEROPLANE FLIGHT CONDITIONS INIT* 00000
ACQUIRE TIME AVERAGED STATIC TEST DATA UNDER
STABILISED POWER CONDITIONS OVER ANGULAR RANGE EMBRACING 10dB-DOWN TIMES OF FLIGHT TIME-HISTORY FOR FACE INDIVIDUAL MEASURED ANGLE* PERFORM 1/3 SPL 1/3xđ Narrowband OCTAVE (& POSSIBLY NARROWBAND) ANALYSIS AND Octave 'ADJUST DATA FOR STATIC PROPAGATION EFFECTS UNDER
REFERENCE AIMOSPHERE CONDITIONS Frequency requency
IDENTIFY ALL RELEVANT DISCRETE TONES AND SUBTRACT FROM 1/3 OCTAVE SPECTRUM TO GIVE BROADBAND SPECTRUM.
CONSIDER TONES AND BROADBAND SEPARATILY
BROADBAND TONES
USING APPROVID SOURCE IDENTIFY ORIGIN OF CONDUCT SOURCE BREAKDOWN IDENTIFICATION TECHNIQUES TONES WHERE APPRO PRIATE Other
ADJUST FOR SOURCE MOTION EFFECTS, ACCOUNTING FOR
RELATIVE VELOCITY, DOPPLER FREQUENCY SHIFT,
CONVECTIVE AMPLIFICATION, DIRECTIVITY EFFECTS, EIC., AND ENGINE THERMODYNAMIC CHANGES
ADJUST FOR NUMBER OF ENGINES, RELEVANT INSTALLA-
TION EFFECTS AND SOURCES OF INTERACTION/ AIRFRAME NOISE, AND RECONSTITUTE 1/3 OCTAVE BAND SOUND
PRESSURE LEVELS EXTRAPOLATE TO AIRCRAFT HEIGHT (SLANT DISTANCE) Airframe
AND ALLOW FOR GROUND REFLECTION, LATERAL PROPAGA-
TION AND OTHER RELEVANT EFFECTS TO PRODUCE QUASI- INFLIGHT 1/3 OCTAVE BAND SPL AND PNLT Aeroplane INTEGRATE PROJECTED PNLT OVER TIME CORRESPONDING PNLT 0000000 TO HEIGHT, FLIGHT PATH AND AIRSPEED 10dB Time Repeat for other angles
COMPUTE HENL EPNI Chap 3 Par 18 31
АС 3б~4B 3/23/88
b. Normalization to Reference Conditions. The analyzed static test data should be normalized to standard conditions in the Part 36 reference atmosphere. The latter correction can only be applied with a knowledge of the spectra of the sources and hence the computations in paragraphs (c), (d), and (e) below should be considered together with the following:
(1) Atmospheric Absorption - Adjustments to account for the acoustical reference day atmospheric absorption are defined in SAE ARP 866A (revised March 15, 1975). The atmospheric absorption should be computed over the actual distance from the effective center of each noise source to each microphone, as described below. In the event that minor differences in absorption values are found in SAE ARP 866A between equations, tables or graphs, the equations should be used.
(2) Ground Reflection - Examples of methods for correcting measured sound pressure levels to standard conditions are described in SAE AIR 1672B-(1983). Spatial distribution of noise sources does not have a first order influence on ground reflection effects and, hence, may be disregarded.
It is also noted that measurements of sound pressure levels with ground-plane microphones may be used to identify the large spectral irregularities caused by interference effects at frequencies less than 1 KHz.
c. Engine Component Noise Source Separation. Unless dominated by just one noise source (e.g., fan noise from a turbofan engine), the measured static engine SPL data at each angular location should be separated into each of the noise sources that may provide significant contributions to the total. This separation should be performed in terms of one-third octave band level and the total noise should be separated into broadband and turbomachinery noise components. To meet the minimum requirement, separation of sources of broadband noise into those due to external jet mixing and those generated internally can be carried out by estimating the jet noise by one or more of the methods identified below, and 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. A further separation of components may be necessary in some instances.
(1) Normally, jet noise dominates the low frequency range of the total noise spectrum and the peak jet noise one-third octave band can readily be discerned. In the mid-frequency range the contribution of jet noise relative to turbomachinery noise sources may not be as apparent. Predictions of jet noise spectral characteristics may be obtained from procedures in SAE ARP 876C, "Gas Turbine Jet Exhaust Noise Prediction," or from other approved methods such as component tests of comparable configurations. These spectral shapes are generally defined as a function of angle and operating condition.
The predicted jet noise spectral shape derived for the relevant position and operating condition should be adjusted to match the measured jet noise level in the low frequency range. The logarithmic sum of the jet and "other" noise sources should be reviewed to ensure that the summation for the mid-frequency range is consistent with the total measured noise. Two other techniques may be useful in obtaining additional information on changes in jet noise levels Chap 3 32 Par 18
(a) Analytical procedures based on correlating full scale engine data with model nozzle characteristics may be useful. Such data have been used to supplement full scale engine data, particularly at low power settings, where noise from other engine sources may make a significant contribution to the broadband noise. Example procedures for coaxial flow jet engines are provided in SAE AIR 1905.
(b) Special noise source locations techniques, such as directional microphones, are available which, when used during full-scale engine tests, can identify the positions and levels of separate engine noise sources.
(2) Turbomachinery noise sources produce both tones and broadband noise. Methods for identifying one-third octave bands containing tones include use of either narrowband analysis or analytical procedures that consider the slopes between adjacent one-third octave bands (analogous to the tone identification procedure in the computation of EPNL).
(3) After separating the total noise into the contributions from each source, the summation, on an energy basis, of the source contributions in each one-third octave band should equal the total noise.
d. Noise Source Position Effects. 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 nozzles and 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 and discrete tones may be considered to be 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:
(1) Spherical Divergence - The distance of the source from the microphone differs from the nominal distance; an inverse square law adjustment should be applied.
(2) Directivity - The angle of the microphone to the source differs from the nominal microphone angle; a linear interpolation should be made to obtain data for the proper angle.
(3) Atmospheric Attenuation - The difference between the true and the nominal distance between the source and the microphone alters the allowance made for atmospheric attenuation.
Chap 3 Par 18 33
e. Engine Flight Conditions. As some thermodynamic conditions within an engine tested statically differ from those that exist in flight, account should be taken of the difference. Noise source strengths may be changed accordingly. Therefore, values for key correlating parameters for component noise source generation should be based on the flight condition and the static data base should be entered at the appropriate correlating parameter value.
Turbo-machinery noise levels should be based on the inflight corrected rotor sprocities that exist at the flight condition. The variation of source noise and jet noise levels should be based on the relative jet levels with key correlating parameters can be determined from the static data base which includes a number of different thermodynamic operating conditions.
f. Noise Source Motion Effects. The effects of motion on jet noise differ from speed effects on other noise sources and, hence, should be considered separately during static-to-flight projection.
1) External Jet Noise. Account should be taken of the frequency-dependent jet relative velocity effects and convective amplificatior effects. Broadly, two sources of information may be used to develop an approved method for defining the effect of flight on external jet noise.
(a) For single-stream engines having circular exhaust geometries, SAE ARP 876C-1985 provides guidance. However, additional supporting evidence may be needed to show when jet noise is the major contributor to the noise from an engine with a more complex nozzle assembly.
(b) Full scale flight data on a similar exhaust geometry 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 supporting information to determine the variation of EPNL, with changes of jet noise spectra at high frequencies.
(2) Doppler Effect. Frequency shifting that results from motion of the source (airplane) relative to a microphone can be accounted for by using the following equation:
f FLIGHT = STATIC (1-M cos 0) where FRATIC - static frequency M = airplane Mach number 0 = angle between the flight path and the sound path connecting the airplane and the microphone It should be noted that for those one-third octave bands dominated by a turbomachinery tone, the Doppler shift may move the tone (and its harmonics) into an adjacent band.
Chap 3 34 Par 18