RC = (33000 P) / W – RS equation (15) where P is engine horsepower, is propeller efficiency, W is airplane weight in pounds, and Rs is airplane power off sink rate in ft/min.
Some aircraft operations handbooks give a power off glide ratio and speed, from which a sink rate can be calculated. The sink rate calculated using these values is not the same as is obtained when operating at best rateof-climb speed and propeller rpm. The best power off g lide condition for an airplane with a variable pitch propeller is obtained with the propeller at low rpm (i.e., a blade pitch angle that provides minimum drag, and usually at an airspeed that is higher than that for best rate-of-climb). Note that in equation (14), if thrust is zero in glide, then the sine of the glide angle is just equal to the ratio of the drag to lift coefficients. For best glide distance this ratio will be lower than the ratio obtained during takeoff climb. For example, for the airplane in Example 1, the aircraft operations handbook states that the airplane will glide 1.7 nautical miles while losing 1000 ft in height, at an airspeed, V, of 105 knots. The glide angle, g, is calculated by observing that its tangent is given by the ratio of the height lost to the distance traveled. Thus:
g = –tan-1[1000 / (1.7 × 6076.1)] = –5.53° equation (16) In this mode of operation C D/CL is equal to the sine of 5.53°, or 0.0964. In the calculation above, it was found that CD/CL was 0.1222 for takeoff with the propeller operated at maximum rpm.
The rate of sink in takeoff configuration can be obtained by conducting glide tests at the speed for best rate-ofclimb with the propeller at high rpm. The time required to lose a fixed altitude, 1000 ft or more, while holding constant airspeed, will give a sink rate that can be used in equation (15). At 90 kt and high rpm, the sink rate for the example airplane is approximately 1125 ft/m.
Propulsive efficiency, that is the product of propeller efficiency and installed power, can be calculated from equation (15) by using this sink rate, in conjunction with the flight manual value for climb rates at a given weight.
Once these factors have been obtained, the climb rate at a different weight can be calculated, and this, coupled with the new climb speed, will allow the climb angle to be computed as in Example 1.
In Example 1 the best rate-of-climb was 970 ft/min at sea level at a weight of 3000 lb. Substituting these values into equation (15), along with the sink rate of 1125 ft/min, gives a product of propeller efficiency and horsepower, for this airplane, of 190. With no change in horsepower or propeller efficiency, substituting these values in equation (15) gives an equation for rate-of-climb at any weight, W:
RC = (6.285 × 106 / W) – 1125 ft/m At a weight of 3200 lb, the rate-of-climb becomes 839 ft/m. At an airspeed of 93 kt, the climb angle is given by:
= sin-1 [839 / (101.3 × 93)] = 5.11° Method 3:
An empirical expression for calculating the rate-of-climb, RC(2), at one weight, W(2), when the rate-of-climb, RC(1), at a different weight, W(1), is known is given by:
RC(2) = RC(1) [W(1) / W(2)]1.5 equation (17) Substituting 3000 lb for W(1), 3200 lb for W(2), and 970 ft/m for RC(1), RC(2) is calculated to be 880 ft/m. In turn, climb angle by this method is given by:
Page 21
= sin-1 [880 / (101.3 × 93)] = 5.36° Reference height for the 3200 lb takeoff weight can now be calculated with each of the three rate-of-climb values by using the equation:
HM = 50 + (8200 – D50M) × tan ( From the rate-of-climb of Method 1:
HM = 50 + (8200 – 2526) tan (5.29°) = 575 ft From the rate-of-climb of Method 2:
HM = 50 + (8200 – 2526) tan (5.11°) = 557 ft From the rate-of-climb of Method 3:
HM = 50 + (8200 – 2526) tan (5.36°) = 582 ft The modest discrepancies among the three calculated values are related to the erroneous assumption that the ratio of propeller efficiency to airspeed is essentially constant. To be conservative, the lower value for HM may be used in the analysis, or alternatively, the average of the three methods, 571 ft.
Step 2. Calculate reference Mach number.
Speed of sound at height h = 571 ft, from equations (3) and (4) is:
cM = 1114.4 ft/s at 571 ft above sea level
b) Air density ratio at 571 ft is given by equations (5-7):
= 0.97954 / 0.99614 = 0.98334
c) Best rate-of-climb speed of 93.0 kt at se a level becomes true airspeed at 571 ft of:
VTAS = 93 / 0.983341/2 = 93.8 kt
d) Airplane Mach number at 93.8 kt is:
MM = (1.688 × 93.8) / 1114.4 = 0.1421
e) Propeller rotational Mach number is given by equation (10):
MRM = 0.8551
f) Helical tip Mach number is given by the square root of the sum of the squares of the airplane and propeller rotational Mach numbers.
MhM = [(0.1421)2 + (0.8551)2]1/2 = 0.8668 Step 3. Reference horsepower remains essentially constant.
Step 4. Calculate the increment in noise level under part 36 Appendix G operating conditions when the airplane is operated at a weight of 3200 lb instead of at 3000 lb, without further changes.
a) Increase in noise level due to decrease in height from 690 ft to 571 ft:
H = 22 log (690 / 571) = 1.81 dB
b) Increase in noise level due to increase in helical Mach number:
From Example 1, Mh = 0.8666.
M = 150 log (0.8668 / 0.8666) = 0.02 dB Note.— Use of “150” for a “k” factor in the tip mach correction is restricted for use only when the Modified mach number is higher than the Parent mach number. Only an approved “k” determined experimentally for the particular propeller airplane combination may be used when the Modified configuration is lower than the Parent.
c) Change in noise level due to change of inflow angle:
In Example 1 the climb angle (γP) was 6.11° with an assumed pitch angle (P) of 4.25°. Using the average climb angle ( γM) obtained from the three methods described above, 5.25° with an assumed pitch angle (M) of 3.00°. Tests conducted by FAA under controlled conditions in a wind tunnel show that for typical small propeller airplane, sound levels increase as climb angle increases due to the change in air-inflow angle to the propeller. On average an increase of 0.5 dB per degree increase of inflow angle was observed in the tests then, the change due to climb angle is given by = 0.5 (M – P) = 0.5 (3.00° – 4.25°) = –0.63 dB equation (18)
d) Total change in noise level:
The total change in noise level, L, is the algebraic sum of the three adjustments.
L = 1.81 + 0.02 – 0.63 = 1.21 dB Clearly, a substantial acoustical change is created if th e weight for this airplane is increased to 3200 lb without any other change to the engine/ propeller installation or other operating conditions. Once it has been determined that an acoustical change has occurred the applicant must develop new noise levels for the modified airplane. This may mean a full part 36 Appendix G noise test mu st be conducted. If the original data used to determine the aircraft noise levels is available to the applicant then additional adjustments might be made (if within limits) to create new noise levels. The noise levels published in the AFM (if available) CANNOT be adjusted using these calculations.
(4.2) Change from fixed to variable pitch propeller If no other change is made to the airplane, such as increasing horsepower, the primary effect of changing from a fixed to variable pitch propeller of the same diameter is a change in Mach number. Typically, fixed pitch propellers are designed to operate optimally under cruise cond itions where the propeller is designed to reach its maximum continuous rated rpm. During climb at the speed for best rate-of-climb the propeller cannot develop anywhere near its maximum rpm. Typically, takeoff rpm is approximately 85 percent of maximum rpm. For example a 150 hp engine having a maximum propeller rpm of 2700 will develop about 2300 rpm at takeoff. A 235 hp engine with a maximum rated rpm of 2575 typically develops about 2200 rpm in takeoff climb.
One of the primary reasons for changing from a fixed to variable pitch propeller is to shorten takeoff distance and to improve takeoff climb rate. An examination of aircraft operations handbooks for airplanes certified with both fixed and variable pitch propellers, at the same horsepower and takeoff weight, indicates the following nominal characteristics:
a) takeoff distances to a height of 50 ft with variable pitch propellers are approximately 88-90 percent of the distance required with a fixed pitch propeller; and
b) at the same takeoff weight and airspeed, the rate-of-climb with a variable pitch propeller is approximately 9-10 percent greater than with a fixed pitch propeller.
The incremental difference in noise level between an airplane equipped with a variable pitch propeller and a fixed pitch propeller can be estimated by using the same adjustment equations used before for height, Mach number, engine power, and climb angle.
Example 2: The aircraft operations handbook for a representative fixed gear airplane with a takeoff weight of 2900 lb has the following performance data with fixed and variable pitch propellers:
Fixed Variable Takeoff distance, ft 1510 1350 Rate-of-climb, 825 900 ft/min Climb speed, Vy, kt 87 87 Propeller diameter, 80 80 in.
Propeller rpm in 2200 2575 climb Using the equations provided in Section 4.1 above, the following calculated quantities are obtained:
Reference height, ft 679 753 Climb angle, 5.37 5.86 degrees Inflow angle (est.), 5.25 6.15 degrees Speed of sound, ft/s 1113.9 1113.6 True airspeed, kt 87.9 88.0 Airplane Mach 0.1332 0.1334 number Rotational Mach 0.6894 0.8072 number Helical Mach 0.7021 0.8181 number The difference in engine horsepower developed with the two propellers in climb can be calculated by using equation (15) to write two equations for the rate-of-climb in terms of horsepower and sink rate, since the sink rate is the same and thus the only difference in rate-of-climb in the two situations is the actual propulsive power. For the fixed pitch case:
825 = (33000Pf) / 2900 – Rs For the variable pitch case:
900 = (33000Pv) / 2900 – Rs Subtracting the first equation from the second results in a difference of about 7 horsepower. The nominal maximum continuous power for the engine is 235.
All the information necessary to calculate the difference in noise levels between the two propeller installations is now available. The difference in the noise levels for the two cases is the algebraic sum of the following adjustments:
a) Decrease in noise level due to increase in reference height:
H = 22 log (679/753) = –0.99 dB
b) Increase in noise level due to increase in helical Mach number:
M = 150 log (0.8181 / 0.7021) = 9.96 dB
c) Increase in noise level due to increase in engine power:
The rated maximum continuous power for the engine is 235 horsepower. The incremental change in noise level is approximately:
P = 17 log [(235+7)/(235)] = 0.22 dB
d) Increase in noise level due to increase in inflow angle:
= 0.5 × (4.15°–3.25°) = 0.45 dB
e) Total change in noise level:
L = –0.99 + 9.96 + 0.22 + 0.45 = 9.64 dB Clearly, an acoustical change exists between the fixed and variable pitch propeller installations unless some other changes are made to the airplane.
(4.3) Change in propeller diameter Calculating the effect on Appendix G noise level of changing propeller diameter without any other change in power or performance is a straightforward application of the following equation:
M= k log (MM/MP) where k equals a constant dependent on the propeller design and Mach number range. A nominal value of 150 is permitted in G36.201 of part 36 if MP is smaller than MM (Parent and Modified respectively).
Although a change in propeller diameter will usually also change performance, if it can be shown that no degradation in takeoff distance or climb performance is created, the airplane performance before the change can be stipulated as the performance after the change. It is clear that an increase in diameter, considered alone, will increase the noise level. It is sometimes desirable to increase propeller diameter to improve climb performance, sometimes at the sacrifice of some cruise capability. The increased noise level caused by the greater diameter can be offset by a reduction in rpm, if the resulting airplane performance is satisfactory. Where this method is used to show "no acoustical change," it will be necessary to pe rform FAA approved takeoff and climb tests to verify Page 25
performance. The following two examples illustrate the e ffect that changes in propeller diameter have on the noise level.
Example 3: The airplane in Examples 1 and 2 is normally equipped with a two-blade propeller having an 84 inch diameter. The same propulsive efficiency is claimed when using the same engine, same propeller rpm, if the propeller is replaced with a propeller of 86 inches diameter. For the same reference height and airplane speed, the difference in noise level between the two installations ma y be calculated as follows. In Examples 1 and 2, the airplane Mach number was 0.1378 and the propeller rotational Mach number was 0.8556 for an 84 inch diameter propeller. For the same altitude and rpm, the rotational Mach number is directly proportional to the propeller diameter, so the rotational Mach number for the 86 inch diameter propeller is 86/84 times 0.8556, or 0.8760. The corresponding Parent and Modified helical Mach numbers before the change and after the change become 0.8666 and 0.8868. The noise level difference is:
M = 150 log (0.8868 / 0.8666) = 1.50 dB If the new propeller has three blades, which would cause a shift in the blade passage frequency, the above analysis would not be adequate to determine the entire effect of the change. The analysis would have to be extended to account for the noise delta caused by the ground reflections and A-weighing filter effects, or flyover tests would have to be conducted.
Example 4: An applicant wishes to replace an existing variable pitch propeller with a diameter of 76 inches for a different pitched propeller with a diameter of 80 inches. The new propeller is designed to provide a somewhat shorter takeoff distance and better rate-of-climb for the airplane. The applicant does not want to go to the trouble of determining the improvement in takeoff distance or climb rate, and is willing to use the existing handbook data to determine reference height.
The applicant has demonstrated to the FAA that, in flight during reference climb conditions, his existing propeller develops 2600 rpm. He believes that, even with a modest limitation on climb rpm to negate the increase in noise level that a larger diameter propeller would normally generate, the larger propeller will still be advantageous. The rpm limitation that would have to be imposed for "no acoustical change" to result, assuming the reference height of 700 ft and true airspeed of 74 kt do not change and the propeller tip shape and thickness are identical, is determined as follows.
Since the airplane climb speed remains the same, the helical Mach number will be the same if the rotational Mach numbers for the two propellers are held constant. The actual reference height and airspeeds are not relevant to this determination. It is not even necessary to calculate the actual Mach numbers, since the rotational Mach numbers will be the same if the products of propeller diameter and rpm remain constant for the two cases. For the original 76 inch propeller, the product is 76 times 2600 or 197,600. The rpm limitation with an 80 inch propeller is therefore 197,600 divided by 80, or 2470 rpm. If this rpm gives satisfactory takeoff and climb performance, limiting climb rpm to 2470 will result in no acoustical change when the larger propeller is installed. Other airworthiness considerations (e.g., engine cooling, stalling, etc.) may need further consideration.
(4.4) Change in engine power Increasing the size of an engine without changing the takeoff weight of an airplane will change the Appendix G noise level. If no change in performance were involved, the level would increase by 17 times the logarithm of the ratio of the increased horsepower to the original horsepower. However, clearly a change in performance does take place. Takeoff distance is shortened, and climb rate is increased. Often, but not always, a propeller change is also made. If the increase in performance offsets the effect on noise level of the change in engine horsepower, no acoustical change occurs. The following example illustrates the required analysis.
Example 5: The airplane in Example 1 is originally equipped with an engine rated at 225 horsepower at 2600 rpm.
The desired installation is for an engine rated at 250 horsepower at 2650 rpm. The same 84 inch two-blade propeller is used in each case. The acoustical effect of this change is evaluated as follows.
Step 1. If propeller efficiency remains the same, takeoff distance varies inversely with engine power and directly as the square of takeoff weight. The takeoff distance to 50 ft height after the increase in horsepower, D50M, without any weight increase, is the original takeoff distance, D 50P, times the ratio of the original horsepower to the increased horsepower:
D50M = D50P × (HPP/HPM) ft In this example:
D50M = 2220 (225) / (250) = 1998 ft Note.— This is an approximation and the final determination for no acoustical change must use FAA approved data from the AFM or the AFMS for the modified airplane. If the applicant chooses not to take credit for (and thus determine per FAA flight test guidelines) the improved takeoff distance, they cannot take credit for the decrease in distance in the acoustical change determination.
Step 2. From equation (15), rate-of-climb for a given airplane configuration is directly proportional to the ratio of engine power to airplane weight, minus the power off sink rate. The constant of proportionality is the unit conversion factor of 33,000 times the propeller efficiency. In Example 1, the product of propeller horsepower developed in climb is less than the rated horsepower due to installation effects and air density less than at sea level. It can be assumed that the same ratio of losses will apply to a slightly higher rated horsepower engine.
Assuming these losses cancel each other out, the apparent propeller efficiency for Example 1 can be stated as the propulsive efficiency, 190, divided by the rated horsepower, 225, for an apparent efficiency of 0.844. Equation
(15) can then be used to determine the rate-of-climb at the increased horsepower:
RCM = (33000) (0.844) (250) / 3000 – 1125 = 1197 ft/m Step 3. Climb angle is given by:
= sin-1 [RC / (1.688 × 60 × Vy)] M = sin-1[1197/(101.3)(90)] = 7.55° Step 4. The reference height after the change, HM, is:
HM = 50 + (8200–1998) tan (7.55) = 872 ft Step 5. Determine reference helical Mach number:
a) At 872 ft the speed of sound is:
cP = 1113.2 ft/s True airspeed is calibrated airspeed divided by the square root of the density ratio:
P= 0.97472
b) Best rate-of-climb speed of 90 kt at sea level becomes a true airspeed at 872 ft of:
VTAS = 90 / 0.97472 1/2 = 91.2 kt
c) Airplane Mach number at 91.2 kt is:
M = (1.688 × 91.2) / 1113.2 = 0.1383
d) Propeller rotational Mach number is:
MR = 0.8725
e) Helical tip Mach number is:
Mh = [(0.1383)2+ (0.8725)2]1/2 = 0.8834 Step 6. Reference horsepower is now 250.
Step 7. Calculate the increment in noise level under part 36 Appendix G operating conditions:
a) Change in noise level due to increase in height:
From Example 1, HP = 690 ft, and from above HM = 872 ft.
H = 22 log (690 / 872) = –2.24 dB
b) Increase in noise level due to increase in helical Mach number:
From Example 2, Mh = 0.8666:
M = 150 log (0.8834 / 0.8666) = 1.25 dB
c) Increase in noise level due to increase in horsepower:
P = 17 log (250 / 225) = 0.78 dB
d) Change in noise level due to change in inflow angle:
In Example 1 the inflow angle was 6.11 degrees with an assumed inflow angle of 4.25°. From equation
(18):
= 0.5 × (4.25°–3.11°) = 0.57 dB
e) Total change in noise level:
The total change in noise level, L, is the algebraic sum of the four adjustments.
L = –2.24 + 1.25 + 0.78 + 0.57 = 0.36 dB The change in noise level due to an increase in horsepower, in this example, is not offset by the effect of improved takeoff performance, resulting in an "acoustical change" of 0.36 decibel. In order to become an NAC situation, some change in rating should take place. A small reduction in allowable takeoff rpm might be reasonable. Several iterations of the above calculations might be necessary to determine the optimum conditions. Any changes in performance where credit is being claimed must be FAA approved, determined using FAA approved methods, Page 28
and published in the FAA approved AFM or AFMS (e.g., improved climb rates, shorter takeoff distances, change in Vy, etc.). Additional airworthiness criteria may need to be reestablished.
(4.5) Increase in drag with no other changes Any modification to an airplane that increases its drag during takeoff and initial climb will increase the Appendix G noise level unless other changes are also introduced. If takeoff distance is increased, or rate-of-climb decreased, or both, the reference height will decrease. The change in level can be calculated from the adjustment equations provided previously if the change in performance can be determined. If an applicant wants to show compliance with the noise regulation via the NAC method, they will probably have to conduct takeoff distance and climb tests. If it looks as though the proposed modification will increase drag sufficiently to make a measurable performance change, then an offsetting change will be required. The amounts of increase or decrease in noise level involved can be calculated by the methods use in the previous examples.
(4.6) Acoustical effects of combined changes Many proposed airplane modifications involve several changes. Examples are replacing an engine and fixed pitch propeller with a higher power engine and a variable pitch propeller, or increasing engine power and takeoff weight at the same time. The acoustical consequences of thes e combined effects can be calculated by combining the methods used in the examples. The sequence for calculating these effects is as follows:
a) Determine the effect changes in takeoff distan ce and rate-of-climb will have on reference height.
b) Determine the effect of helical Mach number caused by any ch anges in airplane speed, propeller rotational speed, or change in speed of sound because of change in reference height.
Note.— The default value of 150 for “k” in the correction equations is only valid when used to determine an increase in helical mach number. It CANNOT be used to reduce the noise level. The 150 factor is a generalized conservative value for simplified calculations where the true “k” value is not determined. Thus it is inherently inaccurate.
c) Determine the effect of an y changes in engine power.
d) Determine any change in noise level due to change in airplane climb angle.
e) Obtain the algebraic sum of the four incremental changes in sound level.
f) If the sum is greater than zero, there are two choices:
— Conduct Appendix G tests as described in the regulation; or — Evaluate different operating conditions to obtain an incremental noise level that is not greater than that of the original airplane. This may be accomplished by derating the engine, reducing fuel/payload, and/or limiting maximum takeoff rpm. One or a combination of these modifications may be iterated until a negative noise increment is obtained by the methods used in the examples.
The cost of not performing an Appendix G noise test should be balanced against the loss of airplane performance or takeoff weight limitations during the lifetime of operation.
101(f) Acquisition of In-duct and/or Near -field Data for Demonstration of NAC
(1) General The FAA has found it acceptable for applicants to conduct engine or component only noise tests to evaluate minor engine changes. Frequently the objective for these tests is to provide evidence that the changes involved produce Page 29
negligible impact on EPNL noise values and may theref ore be categorized as NACs relative to the certificated aircraft configuration. Such testing includes component tests, static engine tests in a test cell, near-field microphone measurements, and in-duct dynamic pressure measurements.
(2) Guiding principles The overall guiding principles to be followed in providing acceptable evidence for substantiation of engine NACs are:
a) the measurements and analyses should adequately model the noise such that small changes in aircraft noise levels can be quantified; and
b) the noise measurement technique and the test e nvironment should not introduce changes to the noise sources that invalidate the predicted small changes in aircraft noise levels.
These guiding principles should be applied in all cases, with details of the approach being justified on a case-bycase basis as appropriate.
Note 1.— It is important that the near-field or in-duct measurements enable a sufficiently accurate prediction of the changes to engine noise in the far field.
Note 2.— It is important that the noise-generating mechanisms of interest are not significantly affected by the test cell environment. The test cell should have an exhaust collector to minimize re-circulation. There should be insignificant inlet distortion or inflow turbulence, or a turbulence control screen or ICD should be employed to minimize such distortions or turbulence. Test cell measurements might not be appropriate for assessing jet noise changes because of the influence of the test cell on the jet development.
Note 3.— Care must be taken to ensure the noise source under investigation is not masked by other unrepresentative noise components. While a reduced acoustic standard of components not under investigation might be acceptable in many cases, there are examples where such differences might invalidate the premise of an NAC (e.g., noise from the intake being masked by a hard-walled bypass duct, or significant noise from an overboard air dump contaminating the measured noise).
(3) Measurement systems Typical measurement systems used to acquire data for substantiation of an engine NAC include:
a) near-field microphones, either in test cells or outdoor facilities;
b) in-duct transducer measurements in the fan inlet or exhaust duct; and
c) core probes to assess combustor or low pressure turbine design changes.
4) Measurement and data analysis procedures The measurement and data analysis process should be accomplished on the basis of the following criteria:
a) an adequate array of transducers should be used to ensure that the measurements adequately model the noise (To determine overall changes in sound pressure level, the measured noise levels will typically need to be averaged azimuthally, radially and/or axially in order to avoid false conclusions being drawn from anomalous readings from single transducers.);
b) environmental conditions in the test setup should be monitored to demonstrate that the local environment
(e.g., test cell temperature) do not cause significant anomalies in the measured noise differences;
c) microphones should be mounted on the test cell wall or on the ground or floor but not in the shadow of any support structures or other test hardware;
d) in-duct transducers should be flush-mounted with minimal loss of area of acoustic treatment. Rakemounted transducers in the flow path should be avoided if they shed wakes that impinge on downstream structures and thereby create significant noise;
e) core probes should be fixed securely to the pylon, boat-tail fairing or other support and not be excessively buffeted by the flow;
f) the specifications of the measur ement system and calibration procedures for microphones, recording and reproducing systems should be in accordance with A36.3 of part 36 (Laboratory calibrations of in-duct transducers and core probes should be conducted be fore and, if possible, after each test. The dynamic range of the transducers should be sufficient to avoid overload.);
g) data should be acquired over the relevant engine operating speeds and for all relevant combinations of engine variables, as specified in the latest version of Reference T13;
h) the interpretation of in-duct meas urements should take into account the possibility that decaying or cutoff acoustic waves may be present that may mask ch anges sensed in the far field of the propagating wave; and
i) two alternative methods could be used in the subsequent analysis of the measured noise levels to demonstrate an NAC:
1) the measured component noise changes could be incorporated into a noise model that predicts the aircraft EPNL. This method has the added value of taking into account in-flight effects and the relative significance of the different noise sources; or
2) in some circumstances, it might be possible to reach the conclusion of an NAC without the need to incorporate the measured component noise changes into a noise model that predicts aircraft EPNL.
The measured noise changes could be examined to see if there is no increase in noise levels at any relevant frequency or engine condition.
Generally, noise models that predict the aircraft noise level expressed in EPNL are based on far-field static test data. Consequently, in either analysis it will be necessary to agree with the FAA on the method for calculating the impact on far-field static noise resulting from near -field microphone measurements or in-duct transducer measurements. This will normally require sound engine ering judgment, seeking out patterns in the data and technical explanations for any observed differences.
Furthermore the statistical uncertainty in the data should be considered. For example if statistical analysis shows that the uncertainty in the data is large and the differences are small, then no conclusions can be drawn from the data. On the other hand if the tests show large decreases in noise levels that outweigh the uncertainty in the data it may be possible to conclude, with reasonable certainty, that the changed engine is indeed quieter than the original engine.
Chapter 2 NOISE STANDARDS: AIRCRAFT TYPE AND AIRWORTHINESS
CHAPTER TABLE OF CONTENTS
201 EXPLANATORY INFORMATION32
201(a) Part 36 - Subpart A - General32
201(a)(1) Applicability and definitions32
201(a)(2) Aircraft configuration34
201(a)(3) Voluntary changes34
201(a)(4) Airworthiness certificates36
201(a)(5) Airworthiness certificates36
201(a)(6) One-Stage limit36
201(a)(7) Stage 1/ Stage 2 limits37
201(a)(8) Date of application37
201(a)(9) Airworthiness compatibility38
201(a)(10) Limitations of part38
201(a)(11) Configuration consistency39
201(a)(12) Dual stage certification39
201(a)(13) Exempt helicopter modifications39
201(a)(14) Helicopter compliance with part 3639
201(a)(15) Noise limits40
201(a)(16) Stage 2 helicopters40
201(b) Subpart B- Transport Category Large Airplanes and Jet Airplanes40
201(b)(1) Noise measurement and evaluation40
201(b)(2) Noise measurement constraints41
201(b)(3) Flight manual statement of Chapter 4 equivalency42
201(c) Subpart H – Helicopters42
201(c)(1) Noise limits42
201(d) Subpart 0 – Operating Limitations; and Information43
201(d)(1) Procedures, noise levels, and other information44
201(d)(2) Non-flight test data50
201(d)(3) Manuals, markings and placards50
201(d)(4) Supplemental information53
201(d)(5) Index of acceptability53
201(d)(6) Airplane limitations53
201(d)(7) Non-complying agricultural and fire fighting airplanes54
201(d)(8) Operating limitations54
201 EXPLANATORY INFORMATION 201(a) Part 36 - Subpart A - General 201(a)(1) Applicability and definitions §36.1(a)
(1) Applicable requirements Page 32