(4) Correction Options: There are two alternatives if the test Mach number is greater than reference. One option is not to make any adjustment, thus ensuring that the noise level is not underestimated.
The second option is to develop approved data, as discussed above, and use the value of k determined in this way.
(5) Correction Calculations: Calculation of the helical tip Mach number requires several steps. The propeller tip speed has two components.
(ii) The speed at which the propeller tip is rotating (e.g., in ft/sec), denoted Vtip
(iii) The forward speed of the airplane (in similar units, e.g., ft/sec), denoted Vtas
The respective Mach numbers for actual test and reference conditions are therefore calculated from:
Vhel = 22 tastip VV + Mhel = cVhel
The speed of sound in air at the test altitude (also similar units, e.g., ft/sec), denoted c.
where Vhel is the helical tip speed of the propeller and Mhel is the helical tip Mach number of the propeller. Calculation of Vtip, Vtas, and c for the appropriate conditions is as follows:
Vtip = rpm x D x π / 60, ft/sec
where rpm is that of the propeller, D is the propeller diameter in feet, and π = 3.14159
Vcas = airplane calibrated air speed multiplied by 1 to convert from knots to ft/sec
s cas tas VV =
where qs ∆=
Pressure Ratio ( )( )2561.568753.61 hE −−=∆
where h is pressure altitude in feet Temperature Ratio ( ) ( )67.45959
67.459 + += Tq , where temperature is in degrees Fahrenheit
The speed of sound c is to be calculated for the appropriate temperature conditions during the test and for the reference condition of 59°F (15°C). The value of c for reference condition calculations can be obtained from
c = , ft/sec at sea level for T = 59°F - lapse rate x h (lapse rate = 0.003566°F per foot)
where h (ft) is the airplane height above mean sea level (MSL) and atmospheric conditions (temperature versus altitude) conform to a standard atmosphere.
MR = Vhel/c day at reference height
MT = Vhel/c test height
c. Procedure
(1) Applicant’s Responsibility: Determine the propeller helical tip Mach number for each noise certification test overflight condition and the reference condition. When required, make the appropriate adjustment to the noise levels to account for the off-reference helical tip Mach number.
414. Section G36.201(d)(4)
Measured sound levels in decibels must be corrected for engine power by algebraically adding an increment equal to
Delta (3) = K3 log (PR/PT)
where PR and PT are the test and reference engine powers respectively obtained from the manifold pressure/torque gauges and engine rpm. The value of K3 shall be determined from approved data from the test airplane. In the absence of flight test data and at the discretion of the Administrator, a value of K3 = 17 may be used.
a. Explanation
This section provides the adjustments for non-reference engine power.
In the engine power correction equation, denoted by Delta (3), PR is the reference engine power setting at HR (airplane over measurement point, calculated for reference flight path), and PT is the actual test power during overflight of the microphone location. These powers are to be determined by referring to the information manual for the engine-propeller configuration and may be assessed from the engine rpm, engine manifold pressure, or other installed instrumentation appropriate to the airplane being tested, at the altitude above mean sea level.
The reference power (PR) will be the maximum continuous installed power for a Vy (best rate of climb) take-off at sea level conditions and standard atmospheric conditions (1013.25 mb, 59°F) adjusted for reference altitude and at 70 percent Relative Humidity. This may be found from the airplane manufacturer's supplied data. If the manufacturer’s data are not available, density ratio for altitude pressure and lapse rate for ambient temperature (59°F – lapse rate x HR, lapse rate = 0.003566°F per foot) calculations can be used.
The actual test power (PT) should be appropriate to the pressure altitude and ambient temperature of the test conditions. Typically, the field elevation plus 1,000 feet can be used to estimate the loss of engine power.
Again, these data should be available in manufacturer’s supplied information for the specific airplane. However, where such information is not available, a correction should be applied to the manufacturer’s published engine power (normally presented for a range of engine speeds under International Standard Atmosphere (ISA) and sea level conditions) to establish the engine power level under the test conditions, as follows:
( ) − = 883.0 117.02 1 s T K KT T TPP , for normally aspirated engines, and
= 2 1 T K KT T TPP , for turbocharged engines
Where PT and PK are the test and known engine powers, TK and TT are the test and known ambient temperatures, and σ is the air density ratio.
Power correction factor K3 can be determined by running a series of tests to obtain noise levels (dB(A)) versus engine power. In cases in which it is not practical to vary engine power independent of engine rpm (such as for fixed-pitch propellers) the power and helical tip Mach number can be obtained from one series of tests and applied to the data.
Section 4.1 of the ICAO Technical Manual (reference 3.j) has guidance on deriving the noise versus tip Mach number and noise versus engine power relationships.
b. Supplemental Information
(1) Fixed-pitch Propellers: This adjustment is not required if the fixed-pitch propeller engine power for the test condition is within 5 percent of the reference engine power.
(2) Maximum Continuous Power for Airplanes with Variable-Pitch Propellers: Airplanes with variable-pitch propellers, in addition to the throttle, have a cockpit control that allows the pilot to control the propeller rotational speed (rpm). Once set, a governor maintains a constant rpm, as long as conditions are within the control range of the governor. In some instances, the engine characteristics are such that the use of take-off power is limited to a specified maximum time (e.g., 5 minutes). Maximum continuous power is less than take-off power, and reducing the engine power setting and/or the propeller rpm sets the lower power. Maximum continuous power settings that are used during the climb portion of the noise certification testing are displayed to the pilot in the form of instrument markings. These markings are in the form of green arcs for ranges where there are no time or other limitations. For piston-engine-powered airplanes, the limiting rpm is usually stated in terms of engine rpm, so with geared propellers the propeller rpm can be determined directly from the gear ratio. The engine power setting parameter, controlled by the throttle, may be the manifold pressure, that is the absolute pressure in the manifold that supplies air to the cylinders. For turboprop installations, there is usually a direct control for the propeller rpm, although it may have only two speeds and may not be infinitely variable. The power setting parameter may be in several different forms, such as propeller (or other) shaft torque, or the temperature of the burned air-fuel mixture at the turbine inlet (TIT).
(3) Maximum Continuous Power for Airplanes with Fixed-Pitch Propellers: Many small airplanes are designed to use a fixed-pitch propeller. With these airplanes, the pilot has only a throttle control with no separate control of propeller pitch. The rpm is dependent on the flight regime and throttle setting. At maximum throttle, as used for take-off and climb conditions, the rpm is appreciably less than maximum, and the engine does not develop its rated power. There is a small increase in rpm as the airplane accelerates on the runway to inflight airspeed, but the rpm is still appreciably below the maximum. Maximum engine rpm cannot be achieved until the airplane is in the cruise phase of flight. The engine power that is achieved during the climb portion of the flight is to be considered the reference power for noise certification purposes and may be dependent on the airport altitude (ambient pressure), temperature, and relative humidity. The airplane manufacturer’s data can provide reference power information.
(4) Determination of Power for Piston Engines: Engine manufacturers provide charts that can be used to determine the power developed by the engine under various operating conditions. The variables that control engine power are rpm, manifold pressure, density altitude, and ambient air temperature. These charts usually apply to a test-bench situation, with no power losses due to power extraction (alternators, pumps, etc.)
and with an air intake and exhaust system designed for test-bench use. In practice, using the un-installed power for both test and reference conditions will provide satisfactory results, when FAA approved.
c. Procedure
(1) Engine Power Adjustment: Determine the engine powers for each test overflight condition and use an approved procedure to adjust the measured noise levels to reference conditions.
415. Section G36.203 Validity of Results
The measuring point must be overflown at least six times. The test results must produce an average noise level (Lamax) value within a 90 percent confidence limit. The average noise level is the arithmetic average of the corrected acoustical measurements for all valid test runs over the measuring point.
(b) The samples must be large enough to establish statistically a 90 percent confidence limit not exceeding –1.5 dB (A). No test results may be omitted from the averaging process unless the omission is approved by the FAA.
a. Explanation
At least six test data points must be obtained in order to satisfy this regulatory section. The certification noise level is then determined from the arithmetic average of these noise levels. All valid data points are to be included in the average. A statistical test is applied to ensure that the average has been adequately determined. The 90 percent confidence limit, computed using established statistical methods, may not exceed ±
1.5 dB (A).
b. Supplemental Information
(1) Valid Data: Test points may not be excluded without the approval of the FAA. Exclusions will not be approved unless there is a valid technical reason. Arbitrary exclusion of a test point on the basis of its measured level is not permitted.
(2) Confidence Limit Exceedance: If the 90 percent confidence limit does not satisfy the ± 1.5 dB (A) requirement, the only recourse is to obtain additional test data points, increasing the number of events until the confidence limit is reduced to ± 1.5 dB (A). The variability of data obtained under controlled conditions should be substantially less than ± 1.5 dB. If the 90 percent confidence interval is near or above the permitted limit, the approved test procedures and/or correction procedures should be carefully reviewed.
c. Procedures
(1) Average Noise Level Calculations: Calculation of the average noise is accomplished by summing the adjusted noise levels of the data points and dividing by the number of values in the sample. This can be written as:
( ) ( ) ( ) ( ) ( )[ ]n n i iavg LALALANLANLA max2max1max1 maxmax ....11 +== ∑=
where LAmax(i) is the corrected noise level of the ith, test flight and N is the total number of test results in the calculation of the average. When the 90 percent confidence limit calculated using data from six or more test flights is within ± 1.5 dB (A), then the average corrected noise level (LAmax)avg resulting from the validated data can be used to determine conformity with the Appendix G Noise Levels, as described in section G36.301.
(2) Confidence Level Calculations: Other statistical terms are used to quantify the scatter of the values around the average value and the range (above and below the average value) that a test result should fall within, with a stated confidence. These terms are known respectively as:
(i) The standard deviation of the results
(ii) The confidence limits (in our case, ±1.5 dB (A)) that are a range above and below the average value for a stated confidence (in our case, 90 percent probability).
The requirement on the confidence limits (of the corrected noise level data) is based on a statistical test involving the number of test results (N) and the standard deviation(s) of the test noise levels.
( ) ( )( )2maxmax111 avgi LALAs −−Ν= ∑Ν=
( ) ( )2111 xxii −−Ν= ∑Ν=
where ( ) == ii LAX max the corrected noise level, in dB (A) of the ith valid test flight ( ) == avgLAx max the average of the corrected noise levels, as described earlier in this section.
After calculating the standard deviation(s), the 90 percent confidence interval can be calculated using the equation:
CI (confidence interval) Ν ±= St where t = Student's t distribution for 90 percent confidence interval and V degrees of freedom (from Table 6)
N = number of data points V = degrees of freedom = N - 1
Table 6: Student’s t Distribution For 90 Percent Confidence Interval
416. PART D - NOISE LIMITS
417. Section G36.301 Aircraft Noise Limits
Degree of Freedom (V) Student's t Distribution for 90 percent Confidence Interval 5 2.015 6 1.943 7 1.895 8 1.860 9 1.833 10 1.812 12 1.782 14 1.761 16 1.746 18 1.734 20 1.725 24 1.711 30 1.697 60 1.671 ∞ 1.645
(a) Compliance with this section must be shown with noise data measured and corrected as prescribed in Parts B and C of this appendix.
(b) The noise level must not exceed 76 dB (A) up to and including aircraft weights of 1,320 pounds (600 kg). For aircraft weights greater than 1,320 pounds, the limit increases from that point with the logarithm of airplane weight at the rate of 9.83 dB (A) per doubling of weight, until the limit of 88 dB (A) is reached, after which the limit is constant up to and including 19,000 pounds (8,618 kg).
Figure G2 shows noise level limits vs. airplane weight.
70.00
72.00
74.00
76.00
78.00
80.00
82.00
84.00
86.00
88.00
90.00 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 Airplane Weight (Pounds x 1,000) Noise Level, dB(A)
Figure 17 - Noise Level vs. Airplane Weight (Figure G2 of part 36, Appendix G)
a. Explanation
The noise level limit varies depending on the maximum certificated take-off gross weight according to the specified formula, as depicted in Figure 17.
b. Supplemental Information
(1) Noise Level Limits: The form of the weight/limit relationship is logarithmic in nature between airplane weights of 1,320 pounds and 3,086 pounds. This format differs from that used to determine the noise level limits for other aircraft under 14 CFR Part 36.
W = Maximum takeoff weight in pounds
0 1320 3086 19000 Noise level in dB (A) 76 32.67 log W – 26.0 88
c. Procedure
(1) Limit Comparison: The noise level limit is computed based on the maximum take-off gross weight and compared with the average measured (and corrected) noise level in terms of dB (A).
418. −429. [RESERVED]
[To be completed later]
[To be completed later]
XVII. 14 CFR PART 36 APPENDIX H- NOISE REQUIREMENTS FOR HELICOPTERS UNDER SUB- PART H [To be completed later] XVIII. 14 CFR PART 36 APPENDIX J- ALTERNATIVE NOISE CERTIFICATION PROCEDURE FOR
HELICOPTERS UNDER SUBPART H HAVING A MAXIMUM CERTIFICATED TAKE-OFF WEIGHT OF NOT MORE THAN 6,000 POUNDS [To be completed later]