# AC 36-4D: NOISE STANDARDS: AIRCRAFT TYPE AND AIRWORTHINESS CERTIFICATION - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Current - Approved: October 12, 2017 - Family: AC 36-4 - Pages: 277 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/632BCB5EDA0858A0862586F40066A8AD.0001 - Page: https://openairdocs.org/ac/36-4 ## Versions - AC 36-4D (Current), approved Oct 12, 2017 - AC 36-4C (Historical), approved Jul 15, 2003 - AC 36-4B (Historical), approved Mar 23, 1988 ## Text ### Page 1 AdvisoryU.S. Department of Transportation Federal Aviation Administration Circular Subject: NOISE STANDARDS: AIRCRAFT Date: 10/12/2017 AC No: 36-4D **TYPE AND AIRWORTHINESS** CERTIFICATION Initiated by: AEE-100 Change: 1. PURPOSE This advisory circular (AC) is intended to promote uniformity of implementation of part 36 of the noise certification requirements of the United States Code of Federal Regulations (CFR) Title 14. It provides (1) guidance to certificating authorities and applicants regarding the intended meaning and stringency of the Standards in part 36 as amended through October 4, 2017 (Amendment 36-31), and (2) the specific pr ocedures that are deemed acceptable in demonstrating compliance with those Standards. Note.— All references in this AC to parts such as part 36, part 21, etc. refer to parts of Title 14 of the United States Code of Federal Regulations. a. Contents: This AC provides FAA policies for implementation of the regulatory language in the form of guidance material and technical procedures relating to the requirements of Appendices A, B, G, J, H and K of part 36 as appropriate. Those appendices describe the noise evaluation methods for compliance with the corresponding chapters of part 36 for jet airplanes, propeller-driven he avy and light airplanes, helicopters, and tiltrotor aircraft. Guidelines are also provided on flight test windows and adjustment of land-use planning data for rotorcraft. The document also includes general information (including administrative information), general and technical references, definitions and bibliographies related to part 36. Information is also provided on equivalent procedures and demonstrating “no acoustical change” for jet airplanes, propeller-driven heavy and light airplanes, and helicopters. b. Scope: This AC provides information for interested parties (such as regulatory authorities, aircraft manufacturers, acoustic consultants, air line and airport operators and others) on compliance with the part 36 noise certification regulations, including FAA policies and interpretations of regulatory language and other technical guidance applicable to noise certification processes for airplane and rotorcraft classes of aircraft. It is applicable for noise certification of normal, utility, acrobatic, commuter and transport category airplanes and normal and transport category rotorcraft as defined in part 21. Information is also provided on flight test windows and adjustment of landuse planning for rotorcraft as developed by the International Civil Aviation Organization (ICAO). This AC does not change regulatory requirements and does not authorize changes in, or deviations from, regulatory requirements. This document contains some mandatory language, where it is appropriate, as well as language that is permissive and advisory in nature. Mandatory language (e.g., the term “must”) is used in the AC for “Explanations of regulatory language”. Mandatory language is also used for certain “Guidance material” and “Technical procedures” to reflect FAA policies regarding compliance with the intent of regulatory language. Permissive language (e.g., the term “should”) is used in this AC, to identify conditions and/or methods and procedures that may be acceptable to the FAA for purposes of aircraft noise certification based upon experience to date, but where an applicant is not constrained from considering or validating and proposing alternatives for compliance with part 36. ### Page 2 c. Document structure: The basic framework of this AC is structured to provide various forms of noise certification guidance material for these aircraft. To a major extent this docum ent follows the structure of the ICAO Environmental Technical Manual (ETM) document 9501 (Reference G5). Chapter 1 provides information on noise certification requirements in part 21. Chapter 2 provides information on FAA regulatory policies. Chapter 3 provides guidance applicable for all aircraft types and subsequent chapters provide guidance unique to different aircraft types. The general format of Chapters 1 through 8 includes one or more of three types of information: explanatory information, equivalent procedures and technical procedures which are described as follows: Explanatory information Explanatory information has the following purpose: a) explains the language of part 36 noise Standards; b) states current policies of the FAA regarding compliance with part 36; and c) provides information on critical issues concerning approval of applicants’ compliance methodology proposals. Explanatory information may take the form of either: a) guidance material (GM) which helps to illustrate the meaning of a specification or requirement; or b) acceptable means of compliance (AMC) which illustra tes a means, but not the only means, by which a requirement specified in part 36, can be met. It may contain reference to an equivalent procedure described in this AC. In the Explanatory sections of Chapters 4 through 8, a heading is provided for each AC section, including a subject or title, and the part 36 appendix and section number to which it refers. For example, GM §A36.2.1.1 is guidance material concerning §2.1.1 of Appendix A. Adjacent to this heading is a bracketed reference to the corresponding section of the ICAO ETM. For this example, “[ETM GM A2 2.1]” is guidance material concerning §2.1 of Appendix 2 of ICAO Annex 16, Volume I (Reference G7). Similarly, AMC §A36.2.1 .1 presents an acceptable means of compliance concerning §2.1.1 of Appendix A and “[ETM AMC A2 2.1]” is an acceptable means of compliance concerning §2.1.1 of Appendix 2 of Annex 16, Volume 1. Equivalent procedures The FAA Office of Environment and Energy (AEE) (located in Washington DC) has approved a substantial number of equivalent procedures proposed by applicants. An equivalent procedure is by definition a test or analysis procedure which, while differing from one specified in part 36, in the technical judgment of the FAA, yields effectively the same noise levels as the specified procedure. Therefore, equivalent procedures provide some flexibility for the applicant in conducting noise certification in accordance with part 36, and may be ap proved for the convenience of an applicant in conducting measurements that are not strictly in accordance with the part 36 procedures, or when a departure from the specifics of part 36 is necessitated by field conditions. Many of these equivalent procedures are referred to in this document because, through their extensive usage by applicants, they have become acceptable practices for aircraft noise certification applications. The acceptability of an equivalent procedure for an aircraft noise certification application is subject to FAA review and approval on a project by project basis. Equivalent procedures such as flight intercept need FAA approval as all other equivalencies, but may not require AEE involvement. ### Page 3 Equivalent procedures fall into two broad categories: a) those which are generally applicable; and b) those which are applicable to a particular aircra ft type; for example, some equivalencies dealing with measurement equipment may be used for all types of aircraft, but a given test procedure may be appropriate only for jet airplanes and not for turboprop airplanes. Typical applications of equivalent procedures requested by applicants are to: a) use previously acquired certification test data for the aircraft type; b) permit and encourage more reliable demonstration of small level differences among derived versions of aircraft; and c) minimize the cost of demonstrating compliance with the requirements of part 36 by keeping aircraft test time, airfield usage, and equipment and personnel costs to a minimum. Technical procedures A technical procedure is a test or analysis procedure not defined in detail in part 36 but which the FAA have approved as being acceptable for compliance with the general provisions of part 36. The procedures described in part 36 must be used unless an equivalent procedure or alternative technical procedure is approved by the FAA. Procedures should not be considered as limited only to those described herein, as this AC will be expanded as new procedures are developed. Also, their presentation does not infer limitation of their application or commitment by the FAA to their further use. Chapter 9 provides guide lines for recertification of airplanes to Stage 4 or St age 5 based upon assessment of those with approved existing noise levels or those specially “modified” to achieve compliance with Stage 4 or Stage 5. Conversion of units Conversions of some non-critical numerical values between U.S. Customary (English) and SI units are shown in the context of acceptable approximations. 2. CANCELLATION Advisory Circular 36-4C, Noise Standards: Aircraft Type and Airworthiness Certification, dated July 15, 2003, is cancelled. 3. REFERENCES General references References that identify FAA Orders, Advisory Circulars, and Reports that provide direction and/or guidance on acceptable processes and procedures for aircraft noise certification are listed in Appendix 1. GENERAL REFERENCES of this AC. These references also include current ICAO noise standards and are noted as Reference G1, Reference G2, etc. in the text. As these documents are referenced in this AC, information is also provided as to the extent that their provisions apply to a given type of aircraft noise certification process. Applicants are encouraged to contact FAA AEE to determine if mo re recent editions are available and app licable to noise certification actions that are being considered for FAA approval. ### Page 4 Technical references References that identify technical documents other than part 36 are listed in Appendix 2. TECHNICAL REFERENCES of this AC. These documents include inte rnational instrumentation standards incorporated by reference in part 36 and are noted as Reference T1, Reference T2, etc. in the text. Bibliography A suggested bibliography relating to certain texts of this AC is provided in Appendix 3. BIBLIOGRAPHY. FAA/DOT addresses Appendix 4. DOT/FAA ADDRESSES lists current addresses of U.S. Department of Transportation and FAA organizations (and noise certification specialists) that might provide assistance to an applicant. 4. DEFINITIONS This section contains two categories of definitions that are used within the regulatory language, policies or other guidance information of this Advisory Circular. The first category defines general terms or organizational designations that depict FAA and applicant roles in the part 36 nois e certification process. The second category defines specific aircraft types that must comply with part 36 regulations and other items related to those aircraft. References to part 1 are to the definitions of 14 CFR Part 1. General definitions: Administrator: The Federal Aviation Administrator or any person to whom he has delegated his authority in the matter concerned. (Ref: part 1) Aircraft certification office (ACO): The (FAA) office that administers the type certificate and production approval of products in the area where the manufacturer is located. Applicant: Section 21.13 of part 21, states: “Any interested person may apply for a type certificate”. The person that applies to the FAA for an aircraft type certificate, whether a new (original), amended, supplemental or provisional type certificate, is identified as the “applicant” and is the entity that must comply with the procedural, certification and production requirements of part 21. Noise certification specialist (NCS): The Directorate (Transport, Small Airplane, or Helicopter) Noise Certification Specialist. Person: An individual, firm, partnership, corporation, company, association, joint-stock association, or government entity. It includes a trustee, receiver, assignee, or similar representative of any of them. (Ref: part 1). Type inspection authorization (TIA): The TIA is the official FAA docu ment that provides direction and communication between the local ACO engineering group and the FAA flight test crew. The TIA usually mandates the provisions of an applicant’s approved demonstration compliance plan and any additional restrictions, inspections, tests, limitations, procedures, etc., that are required by the ACO to be observed during the conduct of the flight or ground testing. Definitions related to aircraft: Acrobatic category airplane: The acrobatic category is limited to airp lanes that have a seating configuration, excluding pilot seats, of nine or less, a maximum certificated takeoff weight of 12,500 pounds or less, and intended for use without restrictions, other than those shown to be necessary as a result of required flight tests. Acrobatic category airplanes are required to show compliance with part 23. ### Page 5 Aircraft: A device that is used or intended to be used for flight in the air. (Ref: part 1) Aircraft engine: An engine that is used, or intended to be used, for propelling aircraft. It includes turbosuperchargers, appurtenances, and accessories necessary for its functioning, but does not include propellers. (Ref: part 1) Airplane: An engine-driven fixed-wing aircraft heavier than air that is supported in flight by the dynamic reaction of the air against its wings. (Ref: part 1) Airport: An area of land or water that is used or intended to be used for the landing and takeoff of aircraft, and includes its buildings and facilities, if any. (Ref: part 1) Civil aircraft: Aircraft other than a public aircraft. (Ref: part 1) Commuter category airplanes: The commuter category is limited to propeller-driven, multiengine airplanes that have a seating configuration, excluding pilot seats, of 19 or less, and a maximum certificated takeoff weight of 19,000 pounds or less, intended for non-acrobatic operations. Commuter category airplanes are required to show compliance with part 23. Helicopter: A rotorcraft that, for its horizontal motion, depends principally on its engine-driven rotors. (Ref: part 1) Large aircraft: Large aircraft means aircraft of more than 12,500 pounds, maximum certificated takeoff weight. (Ref: part 1) Normal category airplane: The normal category is limited to airplanes that have a seating configuration, excluding pilot seats, of nine or less, a maximum certificated ta keoff weight of 12,500 pounds or less, and intended for nonacrobatic operation. Normal category airplanes are required to show compliance with part 23. Normal category rotorcraft: A normal category rotorcraft is a rotorcraft with a maximum weight of 7000 pounds or less and that meets the regulatory requirements of part 27. Primary category aircraft: A primary category aircraft is an aircraft that (1) is unpowered, or is an airplane powered by a single, naturally aspirated engine with a 61-knot or less stall speed, or is a rotorcraft with a 6-pound per square foot main rotor disc loading limitations, and (2) weighs not more than 2700 pounds, and (3) has a maximum seating capacity of not more than four persons, including the pilot, and (4) has an unpressurized cabin. A primary category aircraft is required to meet the regulatory requirements of §21.24 of part 21. Propeller: A device for propelling an aircraft that has blades on an engine-driven shaft and that, when rotated, produces by its action on the air, a thrust approximately perpendicular to its plane of rotation. It includes control components normally supplied by its manufacturer, but does not include main and auxiliary rotors or rotating airfoils of engines. (Ref: part 1). Public aircraft: Aircraft used only in the service of a government, or a political subdivision. It does not include any government-owned aircraft engaged in carrying persons or property for commercial purposes. (Ref: part 1) Restricted aircraft: Used for special purpose operations as defined in §21.25 of part 21. Rotorcraft: A heavier-than-air aircraft that depends principally for its support in flight on the lift generated by one or more rotors (Ref: part 1). Small aircraft: Small aircraft means aircraft of 12,500 pounds or less, maximum certificated takeoff weight. (Ref: part 1) Page 5 ### Page 6 Subsonic airplane: A subsonic airplane means an airplane for which the maximum operating limit speed, MMO, does not exceed a Mach number of 1. Supersonic airplane: A supersonic airplane means an airplane for which the maximum operating limit speed, M MO, exceeds a Mach number of 1. Transport category aircraft: Those aircraft (large or small) that are demonstrated to meet and are certificated to the regulatory requirements of part 25 for transport category airplanes or part 29 for transport category rotorcraft. Jet Airplane: Any fixed wing airplane that is powered by a turbojet or turbofan engine regardless of whether it is an airplane certificated to part 23 (small or commuter aircraft) or part 25 (transport category aircraft). Turbo-propeller airplane: An airplane whose primary source of thrust is from a propeller driven by a gas turbine engine. Utility category airplane: The utility category is limited to airplanes that have a seating configuration, excluding pilot seats, of nine or less, a maximum certificated takeoff weight of 12,500 pounds or less, and intended for limited acrobatic operation. Utility category airplanes are required to show compliance with part 23. 5. ABBREVIATIONS AND ACRONYMS The following abbreviations and acronyms are used within the text of this AC AC = FAA advisory circular ACO = Local FAA aircra ft certification office AEE = FAA Office of En vironment and Energy AFM = Airplane Flight Manual AMC = Acceptable means of compliance APU = Auxiliary power unit C/A = Coarse/Acquisition CAEP = Committee on Aviation Environmental Protection CDI = Course deviation indicator CG = Center of gravity CFR = Code of Federal Regulations CPA = Closest point of approach DAT = Digital audiotape dB = Decibel DER = Designated Engineering Representative DGPS = Differential global positioning system DMU = Distance measuring unit EPNL = Effective perceived noise level ETM = Environmental Technical Manual (“ICAO Environmental Technical Manual. Volume 1, Procedures for the Noise Certification of Aircraft”) FAA = Federal Aviation Administration GDI = Glide-slope deviation indicator GM = Guidance material GPS = Global positioning system Hz = Oscillatory frequency, cycles per second IAS = Indicated airspeed ICD = Inflow control device INS = Inertial navigation system ICAO = International Civil Aviation Organization ILS = Instrument landing system IOD = Issue of data Page 6 ### Page 7 AC 36-40 lRIG lRIGB ISA LAAS LArn ax log LUP MAP MLW MSL MTOW NAC NCS NIST NMEA NPD OAT RFM RH rms rpm SAE SBV SR STC TIA TC TSPI UTC VLA VNTSC VREF 10/12/2017 =Inter-Range Instrumentation Group =Inter-Range Instrumentation Group Standard Serial Time Code Format B =International standard atmosphere = Local area augmentation system = Maximum A-weighted sound level =Logarithmic value to base 10 = Land-use planning = Manifold air pressure = Maximum landing weight = Mean sea level altitude , feet = Maximum takeoff weight = No acoustical change = Noise Certification Specialist =National Institute of Standards and Technology = National Marine Electronics Association = Noise-power-distance = Outside air temperature = Rotorcraft Flight Manual = Relative humidity , % =Root-mean-square = revolutions per minute =Society of Automotive Engineers =Surge bleed valve (engine internal operating valves) =Slant range = Supplemental Type Certificate = Type Inspection Authorization =Type Certificate =Time-space-position information =Coordinated universal time =Very Light Aircraft =Volpe National Transportation System Center (DOT) =Reference Landing Speed u isH Acting Director of Environment and Energy Page 7 ### Page 8 Table of Contents Page 1. PURPOSE … 1 2. CANCELLATION … 3 3. REFERENCES … 3 4. DEFINITIONS … 4 5. ABBREVIATIONS AND ACRONYMS … 6 6-7. RESERVED … 7 **CHAPTER 1. CERTIFICATION PROCEDURES FOR PRODUCT AND PARTS 14 CFR PART 21** 101 EXPLANATORY INFORMATION … 10 CHAPTER 2. NOISE STANDARDS: AIRCRAFT TYPE AND AIRWORTHINESS 201 EXPLANATORY INFORMATION … 32 **CHAPTER 3. TECHNICAL PROCEDURES AP PLICABLE FOR NOISE CERTIFICATION** O F A L L A I R C R A F T T Y P E S 301 NOISE MEASUREMENT PROCEDURES … 58 302 FLIGHT PATH MEASUREMENT … 65 303 ON-BOARD FLIGHT DATA ACQUISITION … 80 304 TIME SYNCHRONIZATION OF MEASURED DATA … 82 305 CALCULATION OF CONFIDENCE INTERVALS … 84 306 ADJUSTMENT OF AIRCRAFT NOISE LEVELS FOR THE EFFECTS … 97 307 NOISE REDUCTION SYSTEMS … 105 308 CALCULATION OF THE SPEED OF SOUND … 107 309 REFERENCE TABLES USED IN THE MANUAL CALCULATION OF … 107 **CHAPTER 4. GUIDELINES FOR TRANSPORT CATEGORY LARGE AIRPLANES,** **SUBSONIC JET AIRPLANES, AND HELICOPTERS EVALUATED UNDER APPENDIX A OR H OF 14 CFR PART 36** 401 EXPLANATORY INFORMATION … 116 402 EQUIVALENT PROCEDURES INFORMATION … 160 403 TECHNICAL PROCEDURES INFORMATION … 192 **CHAPTER 5. GUIDELINES FOR PROP ELLER-DRIVEN SMALL AIRPLANES AND** **PROPELLER-DRIVE COMMUTER CATEGORY AIRPLANES EVALUATED UNDER APPENDIX G OF 14 CFR PART 36** 501 EXPLANATORY INFORMATION … 218 502 EQUIVALENT PROCEDURES INFORMATION … 226 503 TECHNICAL PROCEDURES INFORMATION … 232 CHAPTER 6. GUIDELINES FOR HELI COPTERS NOT EXCEEDING 7000 LB EVALUATED UNDER APPENDIX J OF 14 CFR PART 36 601 EXPLANATORY INFORMATION … 237 602 EQUIVALENT PROCEDURES INFORMATION … 245 CHAPTER 7. GUIDELINES FOR TILTRO TOR AIRCRAFT EVALUATED UNDER APPENDIX K OF 14 CFR PART 3 Page 8 ### Page 9 701 EXPLANATORY INFORMATION … 247 **CHAPTER 8. GUIDELINES ON FLIGHT TEST WINDOWS AND ADJUSTMENT OF** LAND-USE PLANNING NOISE DATA MEASURED IN ACCORDANCE WITH ATTACHMENT H TO ANNEX 16, VOLUME I 801 EXPLANATORY INFORMATION … 253 802 EQUIVALENT PROCEDURES INFORMATION … 260 803 TECHNICAL PROCEDURES INFORMATION … 260 **CHAPTER 9. GUIDELINES FOR AIRCRAFT RECERTIFICATION** 901 INTRODUCTION … 266 902 ASSESSMENT CRITERIA … 267 903 RECERTIFICATION GUIDELINES FOR “MODIFIED”AIRPLANES … 271 Appendix 1. GENERAL REFERENCES … 273 Appendix 2. TECHNICAL REFERENCES … 274 Appendix 3. BIBLIOGRAPHY … 276 Appendix 4. DOT/FAA ADDRESSES … 277 ### Page 10 Chapter 1 CERTIFICATION PROCEDURES FOR PRODUCT AND PARTS 14 CFR PART 21 **CHAPTER TABLE OF CONTENTS** 101 EXPLANATORY INFORMATION … 10 101(a) Part 21 General … 10 101(b) Airworthiness Certification … 10 101(c) Changes in Type Design … 10 101(d) No acoustical changes (NAC) … 11 101(e) Methods for Demonstrating No Acoustical Change (NAC) … 14 101(f) Acquisition of In-duct and/or Near-field Data for Demonstration of NAC … 29 101 EXPLANATORY INFORMATION 101(a) Part 21 General As part of the overall aircraft certification requirements iden tified in part 21, applicants must comply with the noise certification requirements of part 36. Section 21.17 of part 21 specifies that an applicant for a type certificate must show that the aircraft meets the applicable requirements of part 36. In addition, §§21.21 and 21.25 of part 21 specify that an applicant is entitled to a type certificate for an aircraft if the applicant submits the type design, test reports, and computations necessary to show that th e product to be certificated meets the applicable airworthiness and aircraft noise requirements. Similar requirements are imposed for aircraft manufactured in a foreign country for import into the United States (§21.29 of part 21). Refer to the part 21 regulations for specific regulatory language. 101(b) Airworthin ess Certification All aircraft requiring noise certification by the FAA are required to meet the applicable airworthiness requirements of part 21, which identifies the specific airworthiness standards for each aircraft type and product, such as: a) part 23 for normal, utility, acroba tic, and commuter category airplanes; b) part 25 for transport category airplanes; c) part 27 for normal category rotorcraft; d) part 29 for transport category rotorcraft; e) part 33 for aircraft engines; f) part 35 for propellers; g) primary category aircraft; and h) JAR-VLA aircraft. 101(c) Changes in Type Design (1) General Section 21.93(b) of part 21 requires compliance with th e appropriate part 36 “acoustic al change” requirements for a voluntary type design change of an aircraft that ma y increase the noise levels of that aircraft, regardless of whether the change in type design is classified as a minor or major change in §21.93(a) of part 21. ### Page 11 (2) Voluntary changes A general description of FAA handling of various forms of voluntary changes in the type design of an aircraft is addressed in 201(a)(3) of this AC. Procedures for demonstration of compliance with part 21 for specific types of “no acoustical change” (NAC) are presented in 101(d) & (e). 101(d) No acoustical changes (NAC) (1) NAC definition Aircraft/engine model design changes and airframe/engine performance changes may result in very small changes in aircraft noise certification levels that are not acous tically significant. These cha nges are referred to as no acoustical changes (NACs). For this AC NACs, which do not result in modification of an aircraft’s noise certification levels, are defined as: a) changes in airplane noise certification levels approved by the FAA which do not exceed 0.10 dB at any noise measurement point and which an applicant does not track; b) cumulative changes in airplane noise certification le vels approved by the FAA whose sum is greater than 0.10 dB but not more than 0.30 dB at any noise measurement point and for which an applicant has an approved tracking procedure; c) for helicopters certificated according to the standa rds of Appendix H, changes in any one of the noise certification levels approved by the FAA which do not exceed 0.30 EPNdB; and d) for helicopters certificated according to the standa rds of Appendix J, changes in the noise certification level approved by the FAA which does not exceed 0.30 dB(A). (2) Tracking procedures With respect to the tracking procedure referred to in 101(d)(1)(b), noise certification approval has been given based upon the following criteria: a) ownership by the certification applicant of the noise certification database and tracking process on an aircraft/engine model basis; b) when the 0.30 dB cumulative change in the airplane noise certif ication level is exceeded, compliance with the part 36 requirements is required; the airc raft noise certification levels may not be based upon summation of NAC noise increments; c) decreases in noise levels should not be included in the tracking process unless the type design change will be retrofitted to all aircraft in service and included on newly produced aircraft; d) aircraft/engine design changes resulting in nois e level increases should be included in the tracking process regardless of the extent of retrofit to aircraft in service; e) tracking of an aircraft/engine model should, in ad dition to engine design changes, include airframe and performance changes; f) tracked noise increments should be determined on the basis of the most noise-sensitive condition and be applied to all configurations of the aircraft/engine model; g) the tracking should be revised to account for a tracked design change increment that is no longer Page 11 ### Page 12 applicable; h) changes should be tracked to two decimal places (i.e., 0.01 dB). Round-off should not be considered when judging an NAC (e.g., 0.29 dB = NAC; 0.30 dB = NAC; 0.31 dB = acoustical change); and i) an applicant should maintain formal documentation of all NACs approved under a tracking process for an airframe/engine model; the tracking list will be reproduced in each noise certification demonstration document. Due to the applicability dates, some light propeller-driven airplanes are not required to have noise certification levels. However some modifications to these aircraft can be applied which may impact the noise characteristics. In this case, the NAC criteria application should be treated with a procedure approved by the FAA. (3) Criteria for helicopter certification Noise certification approval of modified helicopters should be granted according to the following criteria: a) an NAC approval for a derived version can be made only if the parent “flight datum” helicopter was flight tested to obtain the noise certification levels; b) noise levels for a helicopter designated as an NAC design cannot be used as the “flight datum” for any subsequent design changes; and c) for changes exceeding 0.30 dB, compliance with part 36 requirements may be achieved either by testing or, subject to approval of the FAA, by analytical means; if analy tical means are employed, the noise certification levels cannot be used as the “flight datum” for any subsequent design changes. Note.— Definition for parent “flight datum” aircraft is shown in Table 1. A flow chart illustrating the criteria for dealing with modified helicopters is presented in Figure 1. Due to the applicability dates, some helicopters are not required to have noise certification levels. However some modifications to these helicopters can be applied which may impact the noise character istics. In this case, the NAC criteria application should be treated with a procedure approved by the FAA. (4) Helicopters modifications defined as NACs Changes in helicopter noise level(s) arising from modifi cations associated with the installation or removal of external equipment are defined as NACs and thus need not be determined. For the purposes of this section “external equipment” means any inst rument, mechanism, part, appurtenan ce, or necessary accessory that is attached to, or extends from, the helicopter exterior but is not used, nor is intended to be used, in operating or controlling the helicopter in flight and is not part of an airframe or engine. The following examples are considered to be no acoustical changes: a) the addition or removal of external equipment; b) changes to the airframe made to accommodate the ad dition or removal of external equipment, to provide for an external load attaching means, to facilitate the use of external equipment or external loads, or to facilitate the safe operation of the helicopter with external equipment mounted to, or external loads carried by, the helicopter; c) reconfiguration of the helicopter by the addition or removal of floats and skis; ### Page 13 d) flight with one or more doors and/or windows removed or in an open position; or e) any changes in the operational limitations placed on the helicopter as a consequence of the addition or removal of external equipment, floats, skis, or flight operations with doors and/or windows removed or in an open position. (5) Helicopters with variable noise reduction systems (VNRS) Aircraft can incorporate variable systems, primarily intended to reduce the takeoff/approach noise. Such aircraft can be noise certificated using the guidance provided for aircraft with VNRS. For such changes to existing type designs, the guidelines provided for NAC are applicable. Figure 1. “No-acoustical change” criteria for modifications to noise certificated helicopters Page 13 ### Page 14 101(e) Methods for Demonstrating No Acoustical Change (NAC) for Propeller-Driven Small Airplanes and Commuter Category Airplanes (1) Procedures The methods, as referred to in this section, are aircraft measurement, flight test, analytical or evaluation methods that differ from the methods specified in the text of Appendix G of part 36. The particular method(s) employed must be approved by the FAA. These methods differ from equivalent procedures in that they are used for a finding of NAC under §21.93(b) of part 21 and are not pertinent to part 36. (2) Acoustical changes The regulatory need to evaluate all voluntary changes for acoustical change is found in §21.93(b) of part 21. An acoustical change in the type design of an airplane is defined in §21.93(b) as “…any voluntary change in the type design of an aircraft that may increase the noise levels of that aircraft is an acoustical change…”; note that a change in design that decreases the noise level is not an acoustical change in terms of the rule. There are four categories of small propeller-driven ai rplanes excluded from this evaluation of voluntary changes: 1) Airplanes designated for “agricultural aircraft operations” as defined in §137.3 of part 137; 2); Designated for dispensing firefighting materials; 3) U.S. registered and had flight time prior to January 1, 1955; and 4) Land configured airplanes reconfigured with floats or skis (non-permanent seasonal changes to the landing gear). The exclusions noted in §21.93(b) differ from an earlier definition that applied to propeller-driven small airplanes certificated under Appendix F of part 36. In the earlier definition, acoustical changes were restricted to (1) any change or removal of a muffler or other component of an exhaust system designed for noise control, or (2) any change to an engine or propeller installation which would increase maximum continuous power or propeller tip speed. The current definition of acoustical change in §21.93(b ) for propeller-driven small airplanes and commuter category airplanes is more in line with that for all other aircraft categories. For example, an increase in takeoff weight without any change in engine/propeller installation would not be considered an acoustical change under the earlier definition. Now it would be, since the reference height for noise level measurement would be lower with the increased weight, which would result in a higher noise level. Section 36.9 defines the noise comp liance requirements for propeller-driven small airplanes and commuter category airplanes for which an acoustical change approval is applied for under §21.93(b). Section 36.9 recognizes two airplane categories: those that have previously been shown to comply with part 36, and those that have not. For the first group, any acoustical change must not incr ease noise levels beyond the limits stated in Part D of Appendix G. For the second group, noise levels may not exceed either these same limits, or "the noise level created prior to the change in type design, measured and corrected as prescribed in §36.501 of this part," whichever is greater. This may require that two noise certification tests need be run. The first is to establish the unmodified noise level standard and the second is to measure the modified noise level. Section 36.501(c) also requires that "compliance must be shown with noise le vels as measured and prescr ibed in Parts B and C of Appendix G, or under approved equivalent procedures." The FAA considers any change in type design that will increase the noise level by more than 0.10 dB(A) an acoustical change. However, no acoustical change results as long as any change that can be expected to increase the noise level is offset by other changes in design or performance that will decrease the noise level in such a way that the sum of the changes does not increase the expected noise level by more than 0.10 dB(A). The purpose of this section is to examine methods for calculating changes in noise level introduced by common type design changes and methods for showing compensating noise level effects that will result in a demonstration of NAC. (3) Factors that change noise level Page 14 ### Page 15 Noise levels under Appendix G test procedures are affected by both the basic noise producing properties of the propeller/engine installation and by the takeoff and climb performance of the airplane. Regardless of what mechanical change is made to the airplane, noise levels can be expected to increas e if any of the following conditions results from a type design change (without compensating changes): a) decrease in the height of the airplane at the refe rence distance of 8200 ft (2500 m) from brake release; b) increase in engine horsepower; c) removal or alteration of exhaust mufflers; d) increase in propeller helical tip Mach number; e) increase in propeller tip thickness; f) decrease in airplane climb path angle; g) increase in airplane drag; h) increase in propeller inflow angle; or i) change in the number of propeller blades. Increases expected in the Appendix G noise levels cause d by any of the above changes must be offset by a compensating change to achieve NAC, a nd an appropriate analysis of the ef fects of such a change must be provided. Removal of exhaust mufflers or other changes in the exhaust system may require some form of testing to determine the acoustical effect. Similarly, increasing propeller tip thickness, such as by cutting the tips off an existing propeller to decrease its diamet er, may cause an increase in noise level (which might be offset by the decrease in the propeller rotational Mach number at the same rpm). In both the case of the altered exhaust system and that for the cut-down propeller, a comparison test, during which noise levels are measured for the modified and the unmodified airplanes utilizing part 36 Appendix G flyovers at the same height and airspeed, may be approved as a satisfactory method to determine whether an acoustical change has occurred. Comparison tests are not generally cost saving methods since they may become as complex as a full part 36 Appendix G test. (3.1) Parent aircraft definition Before making any analysis or conducting any test it is necessary to determine what defines the Parent (unmodified) airplane. The Parent airplane represents the noise characteristics existing prior to the change in type design. Table 1 will help identify the Parent aircraft for use in no-acoustical change evaluations. ### Page 16 Table 1. Definition for parent “flight datum” aircraft Modification Parent Aircraft STC/Field Approval/Type Design Change (All changes contained in the same approval). Aircraft defined on the dataplate (Original T.C. configuration) Multiple STC’s/Field Approvals/Type Design Changes (None of which provided noise levels) Aircraft defined on the dataplate (Original T.C. configuration) Multiple STC’s/Field Approvals/Type Design Changes Aircraft defined on the (Where at least one* provides new noise levels and including any dataplate previously approved changes that provided noise levels). (Original T.C. configuration) * The approval that provides new noise values (STC, Service Bulletin, Field Approval, etc.) must be listed in the approval plus the modification that included noise value(s). documentation as required to be previously or concurrently installed. [Should more than one of the included changes contain noise value(s), the approval that provides the lowest value(s), in all applicable reference profiles, will be used. For approvals with more than one reference profile (e.g., Appendix B & H), when one modification does not provide lower, or equal, noise levels in all three profiles, then only one of the previously approved changes may be used as the acoustic parent definition. All other alterations are treated as acoustical changes.] Conversion from one aircraft model into another with no change to the aircraft dataplate. Aircra ft defined on the dataplate (Original T.C. configuration) Conversion from one aircraft model into another with a new or additional dataplate being added. Aircraft defined on the later installed dataplate. (3.2) Definitions Following are definitions used in this section: c = Speed of sound at reference height Page 16 ### Page 17 CD = Airplane coefficient of drag CL = Airplane coefficient of lift D = Diameter of propeller D50 = Distance from brake release to and aircraft height of 50 feet at sea level conditions. H = Reference height over the prescribed measurement location HP = Horsepower F = Force (thrust) M = Mach number RC = Rate of climb at V y rpm = Revolutions per minute T = Standard temperature at reference altitude V = Airspeed Vy = Best rate of climb airspeed. W = Maximum gross takeoff weight  = Propeller inflow angle  = Pressure ratio at reference height. γ = Climb angle θ = Temperature ratio at reference height σ = Density ratio at reference height. Subscripts P = Parent M = Modified R = Rotational A = Airplane TAS = True Airspeed h = Helical tip (4) Examples of type design changes that may affect noise levels Typical applications for type design changes that require an evaluation of the effect of the change on noise levels include: a) increase in maximum takeoff weight without any change in engine or propeller installation; b) change from a fixed pitch propeller to a variable pitch propeller; c) change in propeller diameter; d) change in number of blades; e) increase in engine horsepower; and f) modification that increases drag without any change in engine or propeller installations. Examples include installation of external cargo containers, larger tires on fixed gear, or advertising light arrays. In many cases, changes in noise level introduced by these modifications can be determined analytically by using existing data for the unmodified airplane, or by supplementing the existing data with additional performance Page 17 ### Page 18 information when that performance is approved by the FAA. If that performance is “equal to” or “better” the applicant CANNOT take credit for “better” performance (e.g ., shorter takeoff distance, increased rate of climb, etc.) when evaluating a modification for acoustical change unless the new performance is subsequently published in an approved Airplane Flight Manual Supplement (AFMS) or Supplemental Airplane Flight Manual (SAFM) for the modification. Most requests for supplemental type certificates (STC) are made by applicants that do not have access to manufacturer's data used to determine noise levels, even if these data exist. Since part 36 Appendix G applies to noise certification tests completed after December 1988, much of the existing fleet of small propeller airplanes do not have noise level data evaluated according to part 36 Appendix G. In order to obtain an STC, an applicant must demonstrate an NAC or conduct an entire Appendix G test. Several examples are given here of ways to substantiate a no acoustical change using the adjustment factors described previously. There are, however, some design changes which an applicant may propose that Appendix G compliance is not required, based on similarity to previously approved type certificates (TC) or STCs. Examples of this are older airplanes in a series where the series has several diffe rent engine sizes certified in the same basic airframe. Converting an airplane from one engine power to higher engine power by an applicant for an STC will require the applicant to do a noise analysis, possibly including noise measurements, even though the same engine/airframe combination exists in a production series. The reason for this is that the manufacturer is very unlikely to have demonstrated part 36 Appendix G compliance with any of his production airplanes. Even if the manufacturer has demonstrated compliance and the certificated noise level is published in an FAA Advisory Circular, the data used to demonstrate compliance, including the 90 percent confidence level for the certificated noise level, is proprietary to the manufacturer. Before the adoption of Appendix G, certain changes to an airplane were accepted as an NAC by the FAA without any analysis. Two examples are the replacement of a two- blade propeller with a three-blade propeller, and an increase of less than 2 percent takeoff weight. For the reasons described below, these two types of changes are no longer accepted as an NAC unless the it is substantiated by analysis and/or test data. In the case of propeller replacements with increased number of blades, the change in noise generating mechanisms cannot be simply accounted for by tip Mach number and engine power corr ections. Also, the increase in the number of propeller blades shifts the blade passage to higher frequency, which has a smaller A-weighing filter adjustment. In STC applications with propeller changes, the FAA may require a comprehensive NAC analysis to account for the modified noise generation mechanisms, and the A-weighing filter effects or flyover tests have to be conducted. Increasing the maximum certificated ta keoff weight over the base airplane (or aerodynamic changes which increase drag of the aircraft such that its takeoff performance deteriorates) would cause it to fly at a lower height over the noise measurement location, which results in a higher noise level. An increase in takeoff weight without any change in engine/propeller performance would be an acoustical change unless the noise increase is offset by performance gains or reduction in airplane noise generation, any of which must be substantiated by analysis and/or test data. It is possible to show an NAC condition analy tically as given in Example 1 of this chapter, without actual flight or ground testing. (4.1) Increase in weight For a change in airplane takeoff weight without any change in engine/propeller installation, the following factors influence the noise level under part 36 Appendix G procedures. a) As a reasonable first approximation, the takeoff distance to a height of 50 ft (15 m) is increased by a factor equal to the square of the ratio of the weight after the change to the weight before the change. However, the final determination must be done using the FAA approved takeoff distance from the Approved Airplane Flight Manual (AFM), AFMS, or SAFM. b) The best-rate-of climb speed will in crease essentially as the square root of the ratio of the weight before and after the change. However, the climb speed (and related performance) in the AFM may not be at the Page 18 ### Page 19 best rate of climb speed (Vy) for aircraft type certificated after February 9, 1996. As of this date the climb speed may be dependent on cooling or other airworthiness minimum requirements and not the best rate of climb for the airplane. c) The climb angle at the incr eased weight will be lower. The noise level produced at the higher weight will be greater than at the lower weight because these items combine to generate a lower reference height. The increase in airspeed can also cause a modest change in the reference helical tip Mach number, which may increase noise level. The incremental change in noise level, if nothing else is done, can be shown by an example. Note that the subscript “P” (Parent) is used to denote the "before" conditions and the subscript “M” (Modified) is used to denote the "after" condition. The following examples are provided to give applicants guidance for NAC analysis. The performance calculations used in the examples are for demonstration purposes only; they are not FAA approved performance calculation procedures. Any change in performance data (takeoff or climb) that is to be utilized in the determination of NAC must be published in the aircraft flight manual. Example 1: Calculate the change in noise level expected if th e weight of an airplane increases from 3000 lb to 3200 lb, without any other changes, given the following (note that all of the data given is readily available or is calculable from virtually all aircraft operations handbooks): Step 1 Calculate the parent aircraft reference conditions: From the AFM/AFMS/SAFM/POH: D50= 2220 ft Vy = 90 kt RC = 970 ft/min D = 84 in. rpm= 2600 Calculate: P = sin -1(RC/Vy) = sin -1[970/(60 × /times 1.688 × 90)] = 6.11° equation (1) HP = tan ( D50) + 50 = tan ( 6.11°) (8200 – 2220) + 50 = 690 ft equation (2) DP = 84 in = propeller diameter rpm = 2600 rpm TP = 59(°F) – lapse rate × H P = 59(°F) – 0.003566 × 690 = 56.5°F equation (3) cP = 49.025 × ( T + 459.67) 1/2 = 1113.8 ft/s equation (4) P= (1 – 0.0000068753 × H P)5.2561 = 0.97532 equation (5) P= ( T+459.67) / (59(°F) + 459.67) = 0.99518 equation (6) P= P/P= 0.98004 equation (7) VTAS = V y / P1/2 = 90.9 knots (True airspeed of parent aircraft) equation (8) MPA = V TAS/cP = 1.688 × 90.9 / 1113.8 = 0.1378 (Mach speed of parent aircraft) equation (9) MPR = D P × rpm × π / (60 × 12 × cP) = 0.8556 (Parent rotational Mach number) equation (10) MPh = (M PA2 + MPR2)1/2 = 0.8666 Page 19 ### Page 20 Equation (11) Step 2: Calculate the modified aircraft reference height for a weight of 3200 lb. a) The takeoff distance before the change, D 50P, is 2220 ft. After the weight increase the takeoff distance is given by: D50M = 2220 (3200 / 3000)2 = 2526 ft equation (12) b) The new best rate-of-climb speed at sea level, standard day, V YM, is given by: VYM = 90 (3200 / 3000)1/2= 93.0 kt equation (13) Note.— This is only a first approximation. Final determination of V Y must be done using FAA approved methods. Unless approved and published in the AFMS/SAFM the applicant must use the V Y value published in the original AFM/POH. c) Calculating the climb angle at the increased weight can be done by several methods, depending upon the data available from the aircraft operations handbook or flight manual. Where handbook or flight manual data are not available, data such as climb rate or sink rate may be obtained by performing a limited amount of climb tests. Any such performance tests must be conducted with FAA approval. Method 1: A general equation for calculating climb performance can be obtained if best rate-of-climb data are available in the aircraft operations handbook for different airplane weights. These data can be used to calculate an effective thrust during takeoff, and the effective ratio of drag to lift. For stable climb conditions at moderate climb angles where the cosine of the climb angle may be assumed to be essentially equal to unity, the sine of the climb angle , which is also the ratio of rate-of-climb to climb speed, is determined from: sin (= RC / (101.3 × Vy) = F / W – CD / CL equation (14) Note.— The published climb speed in the AFM/AFMS may not be the actual aircraft V y. The applicant is responsible for determining what VY is. See previous note concerning the appropriate speed. where F is the thrust developed by the propeller, and CD/CL is the ratio of drag to lift coefficients. Thrust is the product of propeller efficiency and engine power, divided by airspeed, with appropriate unit conversions. If it is assumed that the ratio of propeller efficiency to airspeed is approximately constant for airspeeds used for bestrate-of climb, then it can be assumed that thrust remains approximately constant for a given horsepower rating for takeoff. By obtaining RC and V y for two different weights, two simultaneous equations are obtained using equation 14. These may then be solved for F and the ratio of CD/CL, since only the ratio is required. In addition to the data for a weight of 3000 lb, the flight manual gives a rate-of-climb of 1140 ft/min at 2700 lb. The best rate-of-climb speed for a weight of 2700 lb will be the speed for 3000 lb times the square root of the ratio of the two weights, or 85.4 kt. Writing two equatio ns with these values, and by subtracting one from the other, the effective thrust, F, is found to be 686 lb, and the drag/lift coefficient ratio is 0.1222. Substituting the values for thrust, drag/lift ratio, and airspeed for the desired weight of 3200 lb yields climb angle of 5.29°. For a climb airspeed of 93.0 kt, the rate-of-climb is obtained as 868 ft/min. Method 2: Rate of climb, RC, in ft/min, is: ### Page 21 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 ### Page 22 = 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. ### Page 23 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. ### Page 24 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 = (33000Pf) / 2900 – Rs For the variable pitch case: ### Page 25 900 = (33000Pv) / 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 ### Page 26 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. ### Page 27 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: ### Page 28 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 ### Page 29 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 ### Page 30 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; ### Page 31 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. ### Page 32 Chapter 2 NOISE STANDARDS: AIRCRAFT TYPE AND AIRWORTHINESS **CHAPTER TABLE OF CONTENTS** 201 EXPLANATORY INFORMATION … 32 201(a) Part 36 - Subpart A - General … 32 201(a)(1) Applicability and definitions … 32 201(a)(2) Aircraft configuration … 34 201(a)(3) Voluntary changes … 34 201(a)(4) Airworthiness certificates … 36 201(a)(5) Airworthiness certificates … 36 201(a)(6) One-Stage limit … 36 201(a)(7) Stage 1/ Stage 2 limits … 37 201(a)(8) Date of application … 37 201(a)(9) Airworthiness compatibility … 38 201(a)(10) Limitations of part … 38 201(a)(11) Configuration consistency … 39 201(a)(12) Dual stage certification … 39 201(a)(13) Exempt helicopter modifications … 39 201(a)(14) Helicopter compliance with part 36 … 39 201(a)(15) Noise limits … 40 201(a)(16) Stage 2 helicopters … 40 201(b) Subpart B- Transport Category Large Airplanes and Jet Airplanes … 40 201(b)(1) Noise measurement and evaluation … 40 201(b)(2) Noise measurement constraints … 41 201(b)(3) Flight manual statement of Chapter 4 equivalency … 42 201(c) Subpart H – Helicopters … 42 201(c)(1) Noise limits … 42 201(d) Subpart 0 – Operating Limitations; and Information … 43 201(d)(1) Procedures, noise levels, and other information … 44 201(d)(2) Non-flight test data … 50 201(d)(3) Manuals, markings and placards … 50 201(d)(4) Supplemental information … 53 201(d)(5) Index of acceptability … 53 201(d)(6) Airplane limitations … 53 201(d)(7) Non-complying agricultural and fire fighting airplanes … 54 201(d)(8) Operating limitations … 54 201 EXPLANATORY INFORMATION 201(a) Part 36 - Subpart A - General 201(a)(1) Applicability and definitions §36.1(a) (1) Applicable requirements Page 32 ### Page 33 Unless otherwise specified, the date to be used for determining the applicability of the appropriate part 36 standards for an aircraft type design is the date on which application for a type certificate was made. An application for type certification is effective for the time periods specified in §21.17(c) of part 21. (2) Noise related type certification requirements See paragraph 7-8 of Reference G4. (3) Supplemental type certificates See paragraph 7-8 of Reference G4. (4) Definitions of terms Section 4 of the preamble of this AC provides definitions of terms (see Table 2) related to the aircraft types and categories within Items (1) – (4) of §36.1(a). Table 2. Aircraft terminology and categories AIRCRAFT TERMINOLOGY AIRCRAFT TERMINOLOGY Aircraft Large Aircraft Airplane Small Aircraft Jet Airplane Turboprop Airplane (Note 3) Propeller AIRCRAFT CATEGORIES (Note 1) AIRCRAFT CATEGORIES (Note 1) Airplane Helicopter Acrobatic Normal Commuter Transport Normal Primary Restricted (Note 2) Transport Utility Note 1.— Airplane Categories and Aircraft Classes: Part 1 recognizes a distinction between Airplane Categories (e.g., acrobatic , commuter, normal, primary, restricted , transport and utility) and Aircraft Classes (e.g., Conventional Takeoff and Landing (CTOL), Vertical Takeoff and Landing (VTOL), Short Takeoff and Landing (STOL), Reduced Takeoff and Landing (RTO L), and Supersonic Transports (SST)). The noise certification requirements contained in part 36 are specified for several Airplane and Helicopter Categories but not for STOL, RTOL, SST (except for Concorde), or VTOL Aircraft Classes. Note 2.— Restricted Category Propeller-Driven Large Airplanes: Propeller-driven large airplanes that have been type certificated in the restricted category (see §21.25 of part 21) are not required to demonstrate compliance with the requirements of part 36. In the absence of part 36 applicability, an Environmental Assessment is required (see paragraph 7-4 of Reference G4 for additional information). Note 3.— Turboprop Airplanes: Turboprop-powered airplanes are included under the part 36 requirements applicable to propeller-driven airplanes (i.e., subsonic transport category large airplanes; propeller-driven small airplanes, including acrobatic, normal, restricted, and transport categories; or propeller-driven Page 33 ### Page 34 201(a)(2) Aircraft configuration §36.1(b) An applicant’s proposal for an aircraft configuration to be certified under the requirements of part 36 must include any type design changes that may be a result of airworthiness standards testing and analysis, and which would result in acoustical changes. (See §36.1(c) for FAA criteria on when aircraft type design changes represent an acoustical change). If an applicant does not propose the appropriate ai rcraft configuration, additional noise certification testing and analysis may be required. 201(a)(3) Voluntary changes §36.1(c) (1) Examples of voluntary changes in type design A voluntary change in the configuration of an aircraft means a change in the type design of that aircraft. Examples of voluntary changes in aircraft type design that may affect part 36 certificated noise levels are as follows: a) increased engine thrust (powe r) to improve aircraft performance; b) maximum gross weight changes; c) limitations on operational flap settings related to noise; d) hardware revisions (e.g., vortex generators, flap configuration, wing tip stabilizers, wing tip fuel tanks, etc.); e) engine modifications to improve engine performance, accelera tion, deceleration, or surge margin protection; f) addition of wing or body hardware to improve aerodynamics; g) external nacelle changes to improve air flow patterns; h) engine inlet changes to improve ice protection; i) reduction of engine blade tip clearances to improve inlet flow characteristics, and; j) modifications to engine /nacelle acoustic treatments. (2) Acoustical change criteria A voluntary change in the type design of an aircraft that may result in an increase in its certificated noise levels is defined as an “acoustical change.” A type design change that does no t result in an “acoustical change” as defined in section§21.93(b) of part 21 is referred to as a “no acoustical change” (NAC). The numerical criteria used by the FAA in defining an acoustical change for airplanes and helicopters are provided in 101(d) and (e) of this AC and are established to limit the cumulative effects of multiple small changes. (3) NAC cumulative effects When type design changes result in reduced noise levels, an applicant may elect not to provide the full part 36 substantiation, but instead provide FAA with data a nd/or information to subs tantiate an NAC. However subsequent type design changes may require full part 36 certification because of cumulative acoustical effects. Example: An applicant proposes a configuration change which will reduce the noise levels by approximately 2.1 dB for the flyover noise measurement point, 1.1 dB for the lateral noise measurement point, and with no Page 34 ### Page 35 change in noise level at the approach noise measurement point. FAA verifies that the noise levels were reduced, and the applicant accepts the pre-change configuration certificated noise levels as published in the AFM to apply to this change in type design. The applicant may not make another type design change as an NAC that increases the flyover noise level and the lateral noise level without providing substantiation of the cumulative effects of this type design change on AFM noise certificated levels. Multiple type design changes must be evaluated on a cumulative basis (as a combined single configuration) for each noise measurement point for the purposes of determining whether there is an acoustical change. (4) NAC approvals FAA may approve results of component laboratory demonstrations, back to-back noise comparisons from ground tests or flight tests of each configuration, acoustical analyses, etc as acceptable for substantiating that an applicant’s proposed changes to an aircraft type design is an NAC. The results from these evaluations must be applicable to an aircraft’s certificated noise levels at the part 36 reference noise measurement points. (5) Non-standard jet airplane configurations Sections 21.93(b)(2)(i) and (ii) specify exceptions under which certain changes in jet airplane type design are not subject to compliance with part 36 requirements because such configurations are not intended for normal operations, but are typically required for maintenance ope rations (e.g., ferry flights, transporting engines to aircraft that cannot be ferried, etc.). (6) Time-limited changes Time-Limited engine or nacelle chan ges on jet airplanes for periods of up to 90 days (as specified in §21.93(b)(2)(iii) of part 21) may be necessary for continued airplane operation until maintenance can be performed on the original engines or nacelles that were changed. These changes must comply with applicable airworthiness standards and the maintenance, preventa tive maintenance, rebuilding and alteration inspection requirements of part 43. Such changes do not require co mpliance with part 36 requirements even though they may result in an increase in airplane noise levels, relative to the certificated noise levels. FAA policy (see Federal Register Vol. 64, No. 225, Page 65655, dated November 23, 1999) permits the 90 day period allowed by §21.93(b)(2)(iii)of part 21 to continue to be used af ter December 31, 1999, but only for maintenance-related purposes. It cannot be used for meeting Stage 3 requirements due to lack of an adequate number of spare engines or nacelles, or insufficient number of engines or nace lles to operate a Stage 3 fleet at a given time. Section 21.93(b)(2)(iii) may only be us ed to intermix engines when maintenance is required and no conforming engine or nacelle for the configuration is available. (7) Special purpose aircraft Type design changes to propeller-driven small airplanes and helicopters do not need to comply with the provisions of part 36 when they are operated for special purposes as specified in §§21.93(b)(3)(i-iv) and §21.93(b)(4) of part 21. (8) Applicant’s responsibility Pursuant to §21.93(b), an applicant must identify type desi gn changes that affect an aircraft’s certificated noise levels by aircraft serial number and must demonstrat e part 36 compliance for those aircraft. The noise levels approved by the FAA for the modified aircraft must be documented in an FAA Approved AFM or RFM by either a revision to the AFM or RFM (typically an option only available to the aircraft manufacturer), an AFM or RFM Supplement (for Supplemental Type Certificates and Field Approvals), or a Supplemental AFM or RFM (for aircraft that do not have a FAA Approved AFM or RFM). Configurations and modifications for which noise approval is required must be found to comply with applicable airworthiness requirements specified in part 21 as well as the requirements of §36.3 of part 36. ### Page 36 201(a)(4) Airworth iness certificates §36.1(d) (1) Requirements for Issuance of Airworthiness Certificates An airworthiness certificate is issued for each airplane produced, and is a means of distinguishing (by date of issuance) airplanes within a production run that comply with the noise certification requirements of part 36. Prior to the original issuance of a standard airworthiness certificate (see §21.183 of part 21) and regardless of the date of application for certification, compliance with the regulatory requirements of part 36 (including Appendix B) is required for a transport category large airplane, or for a jet airplane regardless of category (see paragraph 7-11(a) of Reference G4). The following table summarizes the requirements of §36.1(d) of part 36. Table 3. Requirements of §36.1(d) of part 36 Maximum Airplane Weight Airplane Type Engine Configuration Effective Date of Part 36 Compliance Date for Airplanes With Initial Flight Time on or After 75,000 pounds or less Subsonic All engines December 1, 1969 December 31, 1974 Greater than 75,000 Subsonic Without Pratt &Whitney JT3D engines December 1, 1969 December 1, 1973 Greater than 75,000 Subsonic With Pratt & Whitney JT3D engines December 1, 1969 December 31, 1974 Concorde – all weights Supersonic All engines October 13, 1977 January 1, 1980 (2) “Grandfathered” airplanes Airplane types not identified above are “grandfathered” airplanes and are permitted to continue normal operations without showing compliance with the requirements of part 36. For example, the twin-engine propeller-driven Convair 580 or Douglas DC-3 airplanes that first logged flight time before December 31, 1974, did not have to comply with the noise provisions of part 36 and may have received a standard airworthiness certificate under the provisions of §21.183 of part 21, or may have received a standa rd airworthiness certificate under the provisions of the regulations in effect at the time of certification without compliance with part 36. 201(a)(5) Airworth iness certificates §36.1(e) (1) Requirements for issuance of airworthiness certificates: An applicant must show compliance of individual airplanes with the requirements of part 36 prior to issuance of airworthiness certificates, as prescribed in §21.183 or 21.185 of part 21 (see paragraph 7-11 of Reference G4). 2) Commuter category airplanes: Since airworthiness standards for commuter category airplanes became effective February 17, 1987, (under part 21, Amendment 21-59) there are no airplanes in this category with initial flight time before January 1, 1980. All commuter category airplanes are thus required to show compliance with the noise certification requirements of part 36. 201(a)(6) One-Stage limit §36.1(g) Page 36 ### Page 37 (1) In-flight configuration changes: In-flight changes in an airplane configuration (including changes in maximum gross weight) that could result in multiple certificated noise levels for a given noise meas urement point or stage classification are not permitted. They are changes in the airplane type design and therefore require a change in operating limitations. A change in maximum gross weight requires an applicant to remove an airplane from service, and comply with the approval procedures of part 43.5 and 43.7. The type design approval procedures under part 21 must be followed if the change does not have prior certification approval. 201(a)(7) Stage 1/ Stage 2 limits §36.1(h) (1) The Stage 1 noise limits apply to all helicopters for whic h application for a type certificate or a change in type design was made prior to March 6, 1986. (2) The Stage 2 noise levels apply to all helicopters for which application for a type certificate or a change in type design is made on or after March 6, 1986, with the following exception. For helicopters for which application for issuance of an original type certificate in the normal, transport or restricted category is made on or after March 6, 1986, and which the FAA finds to be the first civil version of a helicopter which was designed and constructed for, and accepted for operational use by an Armed Force of the U.S. or the U.S. Coast Guard on or before March 6, 1986, it must be shown that the noise levels of the helicopter certificated under the provisions of Appendix H are no greater than Stage 2 levels pl us 2 EPNdB at each measuring point. There is no corresponding provision under Appendix J. The tradeoff provision of H36.305(b) of part 36 may not be used to show compliance with Stage 2 levels plus 2 EPNdB, nor may they be used to increase any noise level beyond these limits. Subsequent civil versions (acoustical changes) must meet the Stage 2 requirements without the extra 2 EPNdB at each measuring point. The tradeoff provisions of H36.305(b) of part 36 may be used in showing stage 2 compliance for these subsequent versions. (3) Helicopters excepted Helicopters that are designated exclusively for agricultural aircraft operations, as defined in §137.3 of part 137, for dispensing firefighting materials, or for carrying external loads, as defined in §133.1(b) of part 133, are excluded from the noise certification requirements. (4) Stage 1/Stage 2 A helicopter, with a maximum takeoff weight of 7000 lb (3175 kg) or less, that fails to comply with the Stage 2 noise limit prescribed in J36.305 of part 36 has the alternative to comply with the Stage 2 noise limit of H36.305 of part 36. Note.— The Appendix J noise limit us es a sound exposure level (SEL) that was defined more stringent by approximately 2 dB than the corresponding average of the three Appendix H noise limits in EPNL. However, this is no guarantee that subsequent Appendix H noise ce rtification testing will comply, especially for a military helicopter seeking civil certification. 201(a)(8) Date of application §36.2 (1) Requirements a) Section 36.2 specifies the applicable part 36 requirements to be those in effect on the date of application for the a) aircraft type certificate or b) change in type design. If the time interval between application for the type certificate (or change in type design ) and the issuance of the type certif icate (or amended or supplemental Page 37 ### Page 38 type certificate) exceeds 5 years, the applicable part 36 requirements are those in effect, on a date to be selected by the applicant, not earlier than 5 years before the date of issuance of the type certificate, amended type certificate, or supplemental type certificate. b) Section 36.2(c) permits the applicant to elect to demonstrate compliance with a later part 36 standard than is required by §36.2(a) or (b). For the purposes of §36.2, “standard” refers to any of the requirements contained in part 36. As specified in §§36.2(c)(1) and (2), an applicant’s election of a later standard is subject to FAA approval, and the election must include standards adopted between the date of application and the effective date of the standard that the applicant has elected to comply with. Thus, an applicant’s election to demonstrate compliance with a particular later standard would also require that compliance be demonstrated with all other part 36 standards (applicable to the aircraft type) ad opted between the date of application and the effective date of the standard that the appli cant has elected to comply with. Furt her, in accordance with §36.2(c)(3), the FAA may require compliance with other standards that were adopted after the date of the elected standard. For the purposes of §36.2(c)(3), “other standards” refers to other part 36 standards. For example, if the compliance standard elected under §36 .2(c)(3) had later been revised, the FAA may determine that it is appropriate for compliance to be demonstrated with the revised standard. Such a determination would also require that compliance be demonstrated with all other part 36 standards (applicable to the aircraft type) adopted between the date of application and the effective date of the standard that the FAA has determined to be the appropriate compliance basis. 201(a)(9) Airworth iness compatibility §36.3 (1) Requirements Aircraft certification tests and analyses must be conducted using configurations, procedures and information that are consistent with the airworthiness re gulations applicable to that aircraft. For example, noise tests or analyses may not be conducted with landing flap settings that could not meet the applicable airworthiness regulation requirements. 201(a)(10) Limitations of part §36.5 (1) Scope of part 36 regulation Part 36 requires compliance with U.S. Government noise standards for aircraft type and airworthiness certification but does not include noise standards related to aircraft operations, such as is provided in part 91. Part 36 certificated noise levels are achieved using procedures that can be duplicated practicably in normal operations by flight crews with safe reserves of thrust (power) and speed. Public Law 103-272 (Reference G2), mandates that the Administrator of the FAA prescribe standards to measure and evaluate aircraft noise. Under this law, an original type certificate may be issued for an aircraft for which substantial noise abatement can be achieved only after the FAA Administrator prescribes standards and regulations that apply to that aircraft. Consideration must be given to relevant information on aircraft noise, consultations with appropriate US and State authorities, whether the standards or regulations are consistent with the highest degree of safety and are economically reasonable, technologically practical, and appropriate for the type of aircraft to which they apply. (2) Compatability with ICAO Annex 16 Standards Part 36 does not contain the Annex allowance to depart from the required reference procedures specified in Annex chapters. For example, §3.6.1.4 of Chapter 3 of Annex 16 permits the certificating authority to approve reference procedures that are different from those contained in §§ 3.6.2 and 3.6.3 of Annex 16, when design characteristics of an airplane would prevent flight from being conducted in accordance with §§ 3.6.2 and 3.6.3 of Annex 16. The FAA recognizes that there may be a need for changes to the specified reference procedures when part 36 may not be appropriate for a particular airplane when design characteristics of an airplane would prevent flight from being Page 38 ### Page 39 conducted in accordance with the reference procedures described in part 36. In cases where part 36 is not appropriate, the rulemaking process, which includes a public comment period, would be followed to develop an appropriate noise certification standard . Thus, the main difference between th e U.S. and ICAO requirements is that under part 36, the reference procedures would be changed through the rulemaking process rather than solely at the discretion of the certificating authority. 201(a)(11) Configuration consistency §36.7(c) (1) Quietest airworthiness approved configuration This requirement was established to show that an acoustical change resulting from an approved type design change must be determined consistently. The FAA policy does not require an applicant to determine the quietest airworthiness approved configuration at the flyover or lateral noise measurement points by noise measurement and evaluation. However, FAA requires consistent evaluations between derivative and originally certificated configurations. For example, an applicant might choose an approved takeoff flap setting of 5 degrees. Later a type design change may be requested that would increase the flyover and lateral noise levels (an acoustical change). The FAA may not require the applicant to measure and evaluate all airworthiness approved takeoff flap settings (e.g., 0, 1, 2, 5, 10 degrees) to determine the quietest configuration, but may, in this example, require appropriate evaluations for a flap setting of 5 degrees. 201(a)(12) Dual stage certification §36.7(e) (1) “Dual-Stage” certificated airplanes Some older airplane models have been certificated to Stage 2 for one configuration and to Stage 3 for another configuration (e.g., airplane takeoff flap setting or maximum allowable gross weights). These airplanes are classified as “Dual-Stage” certificated airplanes. No new dual certifications will be considered. 201(a)(13) Exempt helic opter modifications §36.11 Modifications excluded from compliance with part 36, as listed in part 21, §21.93(b)(4)(ii) “Classification of Changes in Type Design” include: a) addition or removal of external equipment (This means any instrument, mechanism, part, apparatus, appurtenance, or accessory that is att ached to or extends from the helicopter exterior but is not used nor is intended to be used in operating or controlling a helicopter in flight and is not part of an airframe or engine.); b) changes in the airframe made to accommodate the addition or removal to fac ilitate the use of external equipment, to provide for an external load attaching m eans, to facilitate the safe operation of the helicopter with external equipment mounted to, or external loads carried by, the helicopter; c) reconfiguration of the helicopter by the addition or removal of floats and skis; d) flight with one or more doors and/or windows removed or in an open position; or e) any changes in the operational limitations on the helicopter as a consequence of the addition or removal of external equipment, floats, and skis, or flight operations with doors and/or windows removed or in an open position. 201(a)(14) Helicopter compliance with part 36 §36.11(a) Page 39 ### Page 40 (1) Equivalent procedures Subject to prior approval by the FAA, equivalent procedures may be used to measure, evaluate and calculate noise levels (see 1(c) of the preamble of this AC). Some commonly used equivalent procedures relating to helicopter noise certification are included in Chapters 4 and 6 of this AC. (2) Noise level tradeoffs Tradeoffs between the adjusted noise levels for flyover, takeoff and approach defined in H36.305(b) of part 36 may be used to show compliance with the maximum noise levels (noise limits) of Appendix H of part 36. Since Appendix J of part 36 only includes a flyover procedure, no tradeoff is possible. (3) Compliance The procedure used to show compliance under Appendix H or Appendix J of part 36 must follow the applicable parts of the regulation. Compliance with the maximum noise levels (noise limits) prescribed in H36.305 or J36.305 of part 36 must be shown. 201(a)(15) Noise limits §36.11(b) (1) Stage 1 helicopters The maximum noise levels (noise limits) associated with a Stage 1 helicopter certificated under Appendix H of part 36 after a change in type design are dependent on the noise levels of the original Stage 1 helicopter. Unless the Stage 1 helicopter after the change meets the Stage 2 maximum noise levels (noise limits), or Stage 2+ 2 EPNdB levels at each measurement point, it will be necessary to show that the helicopter has no increase in the noise levels at any of the measurement points. This will require determination of the noise levels associated with the original Stage 1 helicopter using Appendix H procedures. If this is not possible an applicant must obtain FAA approval for another procedure. 201(a)(16) Stage 2 helicopters §36.11(c) (1) Helicopter under 7000 lb (3175 kg) If after a change in type design a helicopter previously certificated under Appendix J of part 36, fails to meet the Appendix J Stage 2 maximum noise levels (noise limits) prescribed in J36.305 of part 36 an applicant may attempt to show compliance by electing the use of Appendix H procedures and the Appendix H Stage 2 maximum noise levels (noise limits) specified in H36.305 of part 36. (2) Noise levels Section 36.11 requirements apply to helicopters certificated under both Appendix H and Appendix J of part 36. 201(b) Subpart B- Transport Category Large Airplanes and Jet Airplanes 201(b)(1) Noise measur ement and evaluation §36.101 (1) Appendix A Page 40 ### Page 41 Appendix A prescribes methodologies for conducting airplane noise certification tests and for measurement and evaluation of noise data for comparison with maximum noise levels specified in Appendix B (see Note 1). Specific requirements include: a) specifications of various physical and environmental conditions under which airplane noise certification measurements are permitted; these conditions are related to the noise test site, wind velocity/direction and atmospheric sound attenuation; b) procedures for measuremen t of atmospheric parameters; c) methods for synchronization of airplane position and performance measurements with noise measurements; d) specifications for noise certification measurement and analysis systems (see Note 2); e) methods to calculate atmospheric sound attenuation coefficients and effective perceived noise level (in units of EPNdB); f) methods for adjustment of measured flight test data to reference conditions; g) nomenclature for symbols and units; and h) specification of data to be reported to FAA. Note 1.— Part 36 Appendix A requirements are nearly identical to those of Appendix 2 of Annex 16 Volume 1, Amendment 7. Subsequent changes to Annex 16 Volume 1 after Amendment 7 have not created any significant differences between the requirements of part 36 Appendix A and Annex 16 Volume 1 Appendix 2 at the time of publication of this AC. Note 2.— Some provisions of Appendix A may also apply to the noise certification of helicopters as prescribed in Appendix H of part 36. (2) Equivalent procedures Equivalent Procedures are described in 1(c) of the introductory purpose section of this AC. Section 402 of this AC identifies several potential applications of equivalent procedures for jet and heavy prop airplanes. The ICAO ETM (Reference G5) also provides general and technical guidance on equivalent procedures that have been accepted as a technical means for demonstrating complian ce with the noise certificatio n requirements of ICAO Annex 16, Volume 1. Equivalent procedures provided in the ICAO ETM are incorporated into this AC to the extent applicable to the U.S. regulations.. FAA has not approved these procedures for any current or future aircraft noise certification actions. (See Suppl emental Information Items in Subpart O, §1501, regarding the FAA’s approval process for equivalent procedures). (3) Applicant’s responsibility Applicants must prepare noise compliance demonstration plans for FAA approval that include proposed methods for compliance with the requirements of B36.1 of part 36. Equivalent procedure proposals in these plans must include substantiation of their technical validity and feasibility of application to the airplane type for which noise certification is requested. (See Subpart O §1501, Supplem ental Information, of part 36 for noise compliance demonstration plan documentation requirements.) 201(b)(2) Noise measurement constraints §36.103(a) Page 41 ### Page 42 (1) Noise measurement points The locations of test noise measurement points frequently do not coincide with the reference noise measurement point locations specified in B36.3 of part 36. For example, locations of flyover noise measurement points may be subject to test site anomalies that would make it impractical to set up microphones at the reference locations. An applicant that has an approved equivalent procedure to measure airplane noise data for generation of NPD plots using the Flight Path Intercept procedure may select test noise measurement points (particularly the flyover point) so as to optimize the noise recording system signal-to-noise ratios. (2) Equivalent procedures The FAA does not consider equivalencies that modifies reference flight conditions (see 201(a)(10) paragraph 2 of this AC) without following public process. (3) General approved flight test procedures, principles, methods and flight crew operational practices are included in AC 25-7A, Flight Test Guide (Reference G3). 201(b)(3) Flight manual statement of Chapter 4 or Chapter 5 equivalency §36.105, §36.106 The FAA is not authorized to approve aircraft co mpliance with Annex 16, Volume 1 Noise Standards. (Reference G7). This is the responsibility of the foreign airworthiness authority involved. FAA participation is limited to witnessing tests and reviewing data. For the cases where Stage 3 compliance has been demonstrated it is permissible, but not a requirement, to insert the statement in 201(d) s ection 36.1581(a)(8) of this AC. In those cases where Stage 4 compliance has been demonstrated, the following statement must be included in the AFM or operations manual: “Certification Noise Levels The following noise levels comply with part 36, Appendix B, Stage 4 maximum noise level requirements and were obtained by analysis of approved data from noise tests conducted under the provisions of part 36, Amendment 36-XX (insert part 36 amendment to which the airplane was certificated). The noise measurement and evaluation procedures used to obtain these noise levels are considered by the FAA to be equivalent to the Chapter 4 noise level required by the International Civil Aviation Organization (ICAO) in Annex 16, Volume I, Appendix 2, Amendment 7, effective March 21, 2002.” Similarly, if the compliance demonstration has been demonstarted for Stage 5, the following statement must be included in the AFM or operations manual: “Certification Noise Levels: The following noise levels comply with part 36, Appendix B, Stage 5 maximum noise level requirements and were obtained by analysis of approved data from noise tests conducted under the provisions of part 36, Amendment [insert part 36 amendment number to which the airplane was certificated]. The noise measurem ent and evaluation procedures used to obtain these noise levels are considered by the FAA to be equivalent to the Chapter 14 noise levels required by the International Civil Aviation Organization (ICAO) in Annex 16, Volume 1, Appendix 2, Amendment 11-B, applicable January 1, 2015.” [Incorporated by reference, see § 36.6]. 201(c) Subpart H – Helicopters 201(c)(1) Noise limits §36.805(c) Page 42 ### Page 43 (1) Stage 1 helicopter If §36.805(c) applies, then noise testing will lead to one of the following: a) The helicopter is demonstrated to be a Stage 2 helicopt er: the tradeoff provisions in H36.305(b) of part 36 may be used to demonstrate compliance with the Stage 2 noise level requir ement. The certificated helicopter is not designated as a first civil version, instead the certificated helicopter is designated as a Stage 2 helicopter; b) The helicopter is demonstrated to be a Stage 1 helicopter via the “Stage 2 + 2 EPNdB” noise limits specified in H36.305(a)(1)(ii) of part 36 and is consequently designated as the first civil version. The trade provisions in H36.305(b) of part 36 may not be used for a first civil version; or c) The helicopter fails to demonstrate compliance with the (Stage 2 + 2) noise limits specified in H36.305(a)(1)(ii) of part 36: The helicopter will be denied a type certificate until such time that the helicopter, modified as necessary, successfully demonstrates compliance with at least the requirements for a first civil version outlined above in outcome (b). The helicopter fails to demonstrate compliance if one or more of the demonstrated Appendix H noise levels exceeds the corresponding noise level limit or that flight procedure and the helicopter does not ot herwise meet the Stage 2 limits (note: a helicopter may exceed the Stage 1 limit for a given flight procedure and yet achieve St age 2 compliance through the use of tradeoffs). (2) Change in type design If an applicant has substantiated that a proposed change in type design (as designated in §21.93) would be less noisy or no noisier than the (parent) first civil version helicopter, no additional demonstration of compliance with part 36 is required. The derivative helicopter remains a first civil version. The applicant may elect to: a) carry forward the parent’s noise numbers to the rotorcraft flight manual of the derived version; or b) voluntarily noise test the derivative helicopter in or der to substantiate lower noise certification levels. However, if the derivative helicopter is found to be a Stage 2 helicopter as a result of the voluntary noise test, and if the applicant chooses to certify the derivative helicopter as a Stage 2 helicopter, that derivative helicopter is no longer classified as a first civil version. (3) Subsequent civil version The §36.805(c) requirement for a subsequent civil version applies only to a helicopter previously designated as a first civil version. A subsequent civil version (unless otherwise excepted under §21.93(b)(4) from the acoustical change requirements) is: a) any change in type design of the first civil version that would result in one or more of the noise certification levels of the subsequent civil version that are greater than the corresponding noise certification levels of the first civil version; or b) any change in the first civil version that is suffic ient to require a new type certificate (as provided in §21.19). As specified in §36.805(c), a subsequent civil version must comply with the Stage 2 noise limits. 201(d) Subpart 0 – Operating Li mitations; and Information Page 43 ### Page 44 201(d)(1) Procedures, noise levels, and other information §36.1501 The following types of documentation may be required in implementation of an aircraft (airplane or rotorcraft) noise certification process. (1) Noise certification plans Applicant noise certification plans should be provided to FAA in the early stages of an aircraft program and contain specific types of information including proposals for the methodology an applicant intends to use to certify a given type of aircraft. These plans, when approved, establish an aircraft’s noise certification basis under part 36. When an aircraft program involves agreements between FAA and foreign certificating authorities, these plans may specify an aircraft’s noise certification basis under both part 36 and the noise regulation used by the foreign authority (e.g., ICAO Annex 16, Volume 1). Noise certification plans must include the following types of information: a) an overview of the contents of a forthcoming noise compliance demonstration plan and noise certification report; b) description of applicable noise certification regulati ons and the tests, reports, data submittals and flight manual pages planned to complete an aircraft noise certification process; c) description of conformity requirements; d) DERs and other cognizant personnel to be involved in the aircraft noise certification process; e) milestone schedule of key events; and f) miscellaneous information (e.g., acronyms, organizational codes, etc.). 2) Noise compliance demonstration plans (aircraft flight test) Noise compliance demonstration plans contain the specific methodology (including equivalent procedures) by which an applicant proposes to establis h compliance of a specific aircraft conf iguration with applicable part 36 requirements. These plans must define the information, data, and procedures that an applicant proposes in order to comply with the requirements of part 36. After approval by FAA, flight tests specified in this plan are included as a part of the Type Inspection Authorization (TIA) used to fulfill requirements for TC, STC, and amended TC certification. Noise compliance demonstration plans must include the following types of information: a) introduction: identify applicable part 36 requirements and aircraft noise certification basis, including part 36 amendment level; b) aircraft description: type, model, engines, propellers/rotors, test aircraft serial numbers, engine ratings, engine nacelle acoustic treatments, aircraft gross dimensions (including propeller and/or rotor diameter), airplane engine and ILS antenna locations, aircraft/engine conformity); c) noise certification methodology: test concepts and equivalent procedures, plans for determining airplane engine spooldown characteristics and airframe noise, aircraft/engine performance substantiation, estimated reference conditions including airspeeds, th rust (power) reduction distances, altitudes, thrust (power) settings, methods for determination of EPNL r values and 90% confidence limits, and flight manual format; ### Page 45 d) test description: test site location and characteristics (e.g., topography, ground cover and obstructions), location of noise measurement points, weather constraints, aircraft configuration including flap settings, airbrake and landing gear positions, APU Operation, conditions of pneumatic bleeds and power/takeoffs, test matrix, including weights, target altitudes, airspeeds, engine thrust (power), and test tolerances; e) measurement system components and procedures including calibration procedures : acoustical, meteorological, time/space position, aircraft/engine performance; f) data evaluation procedures including calibration and data processing : approval status of data processing software and version level, adjustment and normalization procedures (acoustical, meteorological, time/space position, and airplane engine performance); g) aircraft certification schedule, including noise certification and TC dates: flight test plans should be submitted to FAA 60 days prior to start of testing. If new and novel equivalent procedures are proposed, or exemptions are required, then submittal of test plans earlier than 60 days prior to the start of testing may be necessary; h) references; i) a listing of all equivalent procedures utilized: for example, flight path intercept procedure; (3) Noise compliance demonstration plans (airplane families) Noise certification of airplane families (derived versions) often require FAA approval of equivalent procedures involving measurement and evaluation of static engine noise test data as described in Chapter 4 of this AC (402(a)(2) for jets and 402(b)(2) for heavy props), in addition to the information listed in Item (2), above. These procedures include projection of static engine noise test data for development of flyover, lateral, and approach NPD plots that define differences between the airplanes used for the original noise certification flight test and a derived version (see Note 1). When use of static engine noise data is proposed, test plans must also be submitted to FAA and either integrated into the basic noise compliance demonstration plan, or submitted to FAA as separate plans and referenced in the basic plan. Static engine noise test plans must include the following information: a) engine description: type, model, test engine serial numbers, nacelle acoustic treatments, engine conformity requirements; b) test description : test facility, engine installation, turbul ence control screen configuration, weather constraints, acoustical and meteorol ogical measurement points, bleed schedules, test run matrix and sequence including thrust (power) settings and test tolerances; c) measurement system components and procedures including calibration procedures : acoustical, meteorological, engine performance; d) data evaluation procedures including calibration, data processing, adjustment and normalization procedures: acoustical, meteorological, engine performance; e) schedules: static noise test, noise certification and TC dates; f) references; and g) a listing of all equivalent procedures utilized. ### Page 46 Note.— Applicants may also generate separate static test plans in cases where development of airplane families is anticipated but not yet formalized and there is opportunity to obtain static engine noise data (because of engine availability) for future noise certification applications. (4) Noise compliance demonstration plans (analytical methods) For certain aircraft type design changes (e.g., weight/thrust, airframe design changes and thrust (power) reduction distances or minor changes in engine components or acoustical treatments), applicants may propose using analytical equivalent procedures to derive noise increments to an aircraft’s certificated noise levels (see 402(a)(2) of this AC for jets and 402(b)(2) of this AC for heavy props) as applicable to the type of aircraft being certificated) or to demonstrate a no acoustical change (NAC) between the original certificated aircraft and the derived version. Noise compliance demonstration plans must include the following information: a) aircraft description: type, model, , engines, propellers/rotors; b) description of type design changes; c) noise certification methodology : overall concepts, description/substantiation of analytical equivalent procedures or methods of assessment of NAC aircraft reference conditions, methods for determination of EPNLr and 90% confidence limit values; d) aircraft certification schedules; and e) references. (5) DER “raw data” reports FAA Order 8110.37E (Reference G6) specifies that Acoustical Designated Engineering Representatives (DERs) may, within the limits of their authority (defined in their letter of appointment) and with prior FAA approval, witness noise certification tests condu cted in accordance with FAA approv ed noise compliance demonstration plans. In cases where tests are witnessed by DERs, the FAA may require the following information to validate that tests are being conducted in accordance with approved plans. a) test event log: includes event times and conditions, nominal engine thrusts (powers) and thrust (power) stability, atmospheric conditions, airspeeds and deviations relative to target speeds, valid and invalid test points (including reasons for invalid points); b) test measurement system components (type/model) and measurement/calibration procedures; c) measurement system calibration events and results; d) measurement system failures, malfunctions, non-standard operations, spurious signals and corrective actions taken; e) report of test condition compliance with part 36 test site requirements; f) field data: samples of measured and corrected noise spectra, acoustical, meteorological, time/space position, aircraft performance field data and data adjustments, EPNdB estimates, and DER notes; and g) summary of meetings. ### Page 47 (6) Noise certification reports Noise certification reports must provide information, data, and procedures demonstrating compliance with the requirements of part 36 and FAA approved noise certification demonstration plans. (7) Noise certification reports (aircraft flight test) These reports must include the following: a) introduction: identify applicable part 36 requirements and aircraft noise certification basis, including part 36 amendment level; b) aircraft description : type, model, engines, propellers/rotors, test aircraft and part serial numbers, aircraft/engine conformity status, reference c onditions (MTOW/MLW, thrust (power), altitudes, airspeeds, takeoff profiles; c) noise certification methodology : noise certification methodology elements of a noise certification demonstration compliance plan approved by FAA for the aircraft configuration that is being certified and the specific report sections in which implementation of each element is addressed; d) test description: test site location and characteristics (e.g., topography, ground cover, and obstructions), location of noise measurement poin ts, test conditions fo r each noise measurement point (including weather), aircraft configuration (e.g., flap, airbrake, landing gear and CG positions, APU operation, conditions of pneumatic bleeds and power takeoffs, airc raft weights, altitudes, airspeeds, engine thrust (power)), engine spooldown measurement points and test conditions, airframe noise measurement points and test conditions, and valid and invalid measurements; e) measurement system components and procedures including calibration procedures : acoustical, meteorological, time/space position, aircraft/engine performance equipment; f) data evaluation procedures including calibration, data processing and adjustment procedures : acoustical, meteorological, time/space position, aircraft/engine performance; g) data analysis and normalization results : analysis results for height of maximum lateral noise, thrust (power) reduction distance, airframe noise adjustment s, engine spooldown characteristics, normalized aircraft data (e.g., noise, power, distance (NPD plots)), values of EPNLr and 90% confidence limits, and aircraft flight manual pages; h) if test witnessing is delegated, then the witness l og or notes specified in section 5 above must be included; and i) references. (8) Noise certification reports (airplane families) Noise certification reports involving airplane family plan concepts must include additional information on noise certification methodology and, if applicable, results of static engine noise tests as follows: a) noise certification methodology: identify approved equivalent procedures for projection of static engine noise data to flight conditions in NPD plot format and methods for defining: NPD plot differences between flight datum (originally certified airplanes) and derived versions, and Page 47 ### Page 48 “Residual” NPD plot differences between flight test data and projected flight data for the originally certified airplane; b) static engine noise test results c) engine description: type, model, serial numbers, nacelle acoustic treatments, engine conformity status] d) test description : test facility, engine installation, turbulence control screen, acoustical and meteorological measurement points, bleed schedules, valid and invalid test points, test conditions for each test point (weather, engine thrust (power); e) measurement systems and procedures including calibration procedur es: acoustical, meteorological and engine performance; f) data evaluation procedures: including calibration and data processing, and adjustment and normalization of acoustical, meteorological, and engine performance; (9) Equivalent procedure applications, documentation and approval process An applicant proposes to use equivalent procedures must be included in aircraft noise compliance demonstration plans (see Note 1). Such proposals may involve new procedures or procedures that have been used previously. However, applicants must be capable of implementing proposed procedures and such procedures must be applicable to the specific aircraft model for which a TC is requested (see Note 2). Equivalent procedures must be substantiated to yield the same noise levels as the procedures prescribed in part 36. FAA Order 8110.4C (Reference G4), paragraph 7-5 specifies AEE as the FAA approving authority for Equivalent Procedures (see Note 3). However, as experience is gained with the application of a particular procedure, AEE may delegate approval of the procedure to ACO or NCS specialists (see Reference G8). For example, FAA policy currently permits NCS approval (or ACO approval when authority is delegated) of NPD equivalencies representing extensions of data used to demonstrate comp liance with part 36 for original certified aircraft. NCS (or ACO) approval authority is limited to cases where either the NPD curve fit is linear and the proposed extension is not greater than 5 percent, or the NPD curve fit is 2 nd order and the proposed extension is not greater than 2.5 percent. The NCS also has approval authority for use of existing NPD plots for amended type designs (excluding extrapolation and increments to NPD plots – see Note 4). Although equivalent procedures may be submitted as proprietary property of an applicant, when several applicants propose closely related or similar procedures (e.g., Flight Path Intercepts), they may be considered to be common industry knowledge and may be described in this AC. Information on equivalent procedures is provided in this AC in Chapter 4, section 402 for subsonic jets, heavy props and helicopters, Chapter 5, section 502 for light props, and Chapter 6, section 602 or light helicopters. Note 1.— Equivalent procedure proposals should be submitted to FAA early in an aircraft noise certification program because approval may require more than one year. Note 2.— Prior to approving an e quivalent procedure, the FAA must assess an applicant’s capability to effectively implement the procedure as well as the appropriateness of the procedure for the specific type of aircraft for which certification is requested. Note 3.— The FAA does not consider equivalencies that modifies reference flight conditions (see 201(a)(10) paragraph 2 of this AC) without following public process. Note 4.— Data obtained during noise certification tests may support extrapolation of noise databases (e.g., for engine thrust (power) range extensions above and below what was tested). Guidance on extrapolation of NPD databases is provided in 402(a)(2)(B)) of this AC. ### Page 49 (10) No acoustical change (NAC) documentation Aircraft type design changes that are categorized as NAC require FAA approval under the provisions of part 21 rather than part 36. Applicants’ proposals for approval of NAC must be submitted to the FAA in writing. The written submittal must define the planned changes in aircraft type design, and must include information, data and analyses that will substantiate that the specified type design change will not result in an “acoustical change” (see Subpart A, §36.1(c)(2) of part 36). NAC substantiation documentation must include: a) introduction (concepts and requirements); b) description of aircraft baseline and proposed type design changes; c) methodology for substantiating an NAC; d) description of tests and/or analyses performed; and e) test and/or analyses results relating to complia nce with FAA NAC criteria (s ee Subpart A, §36.1(c)(3) of part 36). Note 1.— Section 21.93(b) of part 21 states that, “…any voluntary change in the type design of an aircraft that may increase the noise levels of that aircraft is an ‘acoustical change’.” Therefore NAC is determined on an aircraft by aircraft basis rather than just on the basis of the type certificate data sheet (TCDS) information for the aircraft model. Note 2.— Section 183.29(i) of part 183 specifies that Acoustical DERs may not determine that a type design change is an NAC. As such, Acoustical DERs may not make recommendations for approval of NACs under Form 8110-3. (11) Other noise-related documentation FAA develops noise-related documentation other than that required of applicants under Subpart O, §1501 of part 36 as follows: a) Findings required by public law 103-272 (Reference G2): Compliance with Public Law 103-272 requires that FAA, before issuing an original TC for any aircraft of any category, regardless of whether part 36 applies to the aircraft, must make a finding to determine whether: Substantial noise abatement cannot be achieved for that aircraft by prescr ibing standards and regulations consistent with the limitations of section 44715(b), or Substantial noise abatement may be so achieved in which case the regulatory process must be used to determine the extent of noise reduction to be required before an original TC may be issued. Note.— The specific process required for FAA to complete Noise Finding documentation is presented in paragraph 7-3 of FAA Order 8110.4C (Reference G4). b) Environmental Assessments: Compliance with the National Environmental Policy Act of 1969 may require that the FAA conduct an Page 49 ### Page 50 Environmental Assessment (EA) as specified in Ch apter 4 of FAA Order 1050.1F (Reference G1 xx). An EA is conducted when type certificate actions (New, Amended, or Supplemental) are initiated for aircraft types for which part 36 requirements do not exist (see Note 1). An Environmental Impact Statement (EIS) or Finding of No Significant Impact (FONSI) must be prepared by the FAA (see Note 2). The FAA may grant an interim aircraft noise certification upon completion of an EA while appropriate rulemaking is being completed. Note 1.— Applicants may be requested to provide appropriate noise data to FAA to support the EA and/or rulemaking processes Note 2.— Procedures for generation of an EA, EIS or FONSI are also found in FAA Order 1050.1E (Reference G1). (12) Procedures a) Applicant’s responsibility : Applicants must generate and submit to FAA the following noise certification documentation:  Noise certification plans;  Noise compliance demonstration plans (including proposed equivalent procedures);  Noise certification reports;  NAC substantiation. b) DER’s responsibility : Acoustical DERs responsibilities are as defined in FAA Order 8110.37E (Reference G6). 201(d)(2) Non-flight test data §36.1501(b) (1) Airplane families Chapter 4 of this AC, 402(a)(3) for jets, and 402(b)(3) for heavy props, present guidance on static engine noise tests and projection of static engine noise data to flight conditions. Comparisons of projections of static engine noise data from the engines of an originally certified airplane with those of modified configuration engines could enable certification of modified configurations without further expensive and time consuming flight tests (see §36.1501(b) of part 36). An originally certified airplane and its derived versions are known as an “Airplane Family”. (2) Tests for supplemental data: An applicant should attempt to obtain static engine or flig ht test noise data at engine thrust (power) settings and ranges, and airplane flight configurations and conditions that may be needed in the foreseeable future. This may necessitate over-boosting of the engine thrust (power) (with the permission of the engine manufacturer), resetting of the engine idle thrust (power) to lower limits, testing to higher or lower range of airspeed and angles of attack, testing at higher or lower altitudes, etc. 201(d)(3) Manuals, markings and placards §36.1581(a) (1) Aircraft flight manual Page 50 ### Page 51 An approved Airplane Flight Manual (AFM) or Rotorcra ft Flight Manual (RFM) is required for certification of each aircraft type in compliance with §36.1581 of the Airworthiness Standards for parts 21.24 (Primary Category), 23, 25, 27, or 29. Section 36.1581 of part 36 requires that certificated noise level compliance information be included in an AFM or RFM. These levels must be cont ained in an FAA approved section of an AFM or RFM other than the limitations section. Fo r example, the performance section is an appropriate place to include the certificated noise level compliance inform ation. The certificated noise levels ar e to be reported to 0.1 dB in the AFM or RFM. Further, an AFM or RFM must specify the ty pe certificate limitations if any, that are established as a result of part 36 compliance, combined with the airworthiness limitations. (2) Rotorcraft flight manual In addition to the requirements of §36.1581, §H36.3(e) and §J36.3(c)(1) of part 36 require that the values of V H and VNE defined for noise certification must be listed in the RFM. The value of V NE is located in the limitations section of the RFM. The value of VH should be listed in the section of the RFM containing the certificated noise level compliance information. (3) Airplane flight manual limitation section a) An AFM may be issued for a single airplane or a group of similar airplanes (including more than one series of a specific model). The AFM must clearly identify (by airplane serial number) the operating limitations, including the maximum weight limits that apply to an airplane. b) The AFM may address a single configuration (har dware build) or multiple configurations. If an AFM includes information for more than one configuration (hardware build), the appropriate airplane operating limitations must be clearly identified for each config uration. Furthermore, if not all configurations are approved for the airplanes listed in the AFM, the AFM must clearly identify by se rial number, the proper operating limitations for each airplane. c) The operating limitations contained in the Limitations Section (including any noise-limited weights) must be expressed in mandatory language, not permissive language. The terminology used in the AFM must be consistent with the relevant regulatory language. (4) Multiple airplane noise certifications a) Multiple noise-limited gross weight pairs (takeoff and landing) for one airplane configuration are not permitted by Subpart A, §36.1(g) for compliance with part 36. Only one set of gross weight limits that pertain to a particular configuration (hardware build) may be established under part 36 for a particular airplane. b) An airplane configuration (hardware build) may be certificated only to a single Stage as it qualifies under part 36 (Stage 1, Stage 2, Stage 3, Stage 4, or Stage 5). The limitations section of the AFM for an airplane configuration may not contain operating limitations that would result in noise levels exceeding part 36 noise limits. However, a Stage, 2 airplane may be recertif icated as Stage 3, provid ed that it meets the noise limits for Stage 3 and the AFMs are revised and approved for configurations appropriate for compliance with Stage 3, and the Stage 2 approval is deleted. (5) Landing flap restriction a) An operating limitation preventing the use of an approved landing flap setting cannot be imposed under part 36 and must be established under the airworthiness requirements or through a voluntary type design change. If such a restriction is requested by an applicant in order to comply with part 36 certification requirements that flap setting limit must be incorporated into the design and operation of the airplane. b) For some Stage 3 airplanes, a “softguard” (such as a crushable cover plate) or frangible device must be Page 51 ### Page 52 installed on the flap selection control to safeguard its use in instances other than emergency. This device makes obvious the flap settings that are not to be used for normal operation. Fracture or deformation of the softguard would indicate use of landing flap settings that have not been approved in demonstrating compliance with part 36. For airworthiness purposes, the design of these devices allows the flight crew use of airworthiness approved restricted flaps in emergencies (declared and undeclared). The device may not be resettable by the flight crew; a damaged softguard or frangible device must be repaired as a maintenance action accomplished in accordance with part 43 prior to the next flight. (6) Procedures a) Applicant responsibility: An applicant must identify limiting configurations that may be required to satisfy airworthiness and/or noise regulatory requirements. The limiting configurations, including approved aircraft gross weights, are identified in the “Limitation Sections” of the approved aircraft AFM or RFM. (7) Implementation of landing flap restrictions FAA policy allows a choice of two options for an applicant to implement landing flap restrictions, where their approval is consistent with applicable airworthiness requirements: a) where permitted and possible, remove the cockpit flap selection control and performance information from the Performance Section of the original AFM that is relevant to flap settings not considered in demonstrating compliance with part 36; and b) if flap settings not considered in demonstrating comp liance with part 36 are to remain selectable then, to prevent their use in normal operations: — Restrictions against their use, except under emer gency (declared and undeclared) operations, must be incorporated into the AFM Limitations Section. — Operational information must be relocated to the Emergency Procedures Section. — For Stage 3 airplanes, must install a “softguard” or frangible device over the flap selection control. — Placards may be installed to provide appropriate information to the flight crew. (8) ICAO certifications The FAA is not authorized to approve aircraft compliance with Annex 16, Volume 1 Noise Standards. (Reference G7). This is the responsibility of the foreign airworthin ess authority involved. FAA participation is limited to witnessing tests and reviewing data. If the Stage 4 or Stage 5 compliance has been demonstrated refer to §36.105 or §36.106 of part 36. In those cases where Stage 3 compliance has been demonstrated, it is permissible to insert the following statement in the AFM: “Certification Noise Levels The following noise levels comply with part 36, Appendix B, Stage 3 noise level requirements and were obtained by analysis of approved data from noise tests conducted under the provisions of part 36, Amendment 36-XX (Insert part 36 amendment to which airplane was certificated). The noise measurement and evaluation procedures used to obtain these noise levels are essentially equivalent to those required by the International Civil Aviation Organization (ICAO) in Annex 16, Volume I, Amendment 7, Chapter 3. ICAO Annex 16, Volume I, Chapter 3 approval is applicable only after endorsement by the Civil Aviation Authority of the country of airplane registration.” ### Page 53 201(d)(4) Supplementa l information §36.1581(b) (1) Publication The FAA allows publication of supplemental noise information in an approved AFM or RFM for configurations other than the maximum takeoff and landing weight configuration or noisiest approach configuration. Such information may be useful to airlines in conjunction with their selection of operational route structures and airport operational requirements. (2) Applicant’s responsibility Applicants must identify supplemental noise levels as “Supplemental Information” when they are included in an AFM or RFM. Even though noise data used to provide “Supplemental Information” may have been obtained during certification noise testing and witnessed by the FAA, it must still be clearly marked to show that it is not intended to demonstrate compliance with part 36, and may be presented in a non-FAA approved section or page of the AFM or RFM. 201(d)(5) Index of acceptability §36.1581(c) (1) Certification vs. operational rule Part 36 is a certification rule and not an operational ru le. The certificated noise levels may or may not meet the operational noise requirements for any particular airport or operating rule. The airc raft noise certification provisions of part 36 do not place any operational permissi bility or limitations on th e operation of certificated aircraft at any airport. (2) Applicant’s responsibility Applicants are responsible for developing an AFM or RFM and obtaining FAA approval prior to the aircraft entering service. The approved flight manual must list the certificated noise levels that have been shown to comply with part 36 maximum noise levels (noise limits) and contain the following statement as specified in §36.1581(c) and reprinted here: “No determination has been made by the Federal Aviation Administration that the noise levels of this aircraft are or should be acceptable or unacceptable for operation at, into, or out of, any airport.” 201(d)(6) Airplane limitations §36.1581(d) (1) Airworthiness maximum gross weights Transport category large airplanes and jet airplane gross weights are normally limited for structural, performance or economic reasons and demonstrated to the appropri ate airworthiness standards for the airplane. Those airworthiness limited gross weights may constitute the maximum operational gross weights and are to be identified in the “Limitations Section” of an approved AFM. (2) Approach noise limiting configurations Applicants may need to limit the landing flaps or the approach gross weight in order to comply with the requirements of part 36. All noise limiting gross weights and configurations are to be included in the “Limitations Section” of an approved AFM. ### Page 54 (3) Takeoff noise limiting configurations An applicant may establish takeoff limiting configurations (e.g., maximum takeoff gross weight, takeoff airspeed schedules, takeoff thrust (power) derate schedule, in-flig ht APU operation) that are more restrictive than the airworthiness limited configurations. These noise limited configurations must be furnished in the Limitations Section of an approved AFM. (4) Optional engine thrust (power) ratings (derate and reduced thrust) Compliance with part 36 is only required at full rated takeoff thrust (power). However, an airplane may be type certificated at derated/reduced thrust (power) that is less than full rated takeoff thrust (power). This is not an acoustical change as defined in §21.93(b) of part 21 provided that the full rated thrust (power) remains approved for a given airplane configuration. Airplane type certification at derated/reduced thrust (power) does not prohibit an applicant from employing thrust (power) reduction for noise certification as permitted by B36.7(b) of part 36. 201(d)(7) Non-complying agricultural and firefighting airplanes §36.1583(a) (1) Fire fighting airplanes Those small propeller-driven airplanes that are specifically designed for “dispensing of fire fighting materials" are exempt from the noise certification requirements of part 36. (2) Agricultural airplane operations Agricultural aircraft operation, as defined in §137.3 of part 137, means the operation of an aircraft for the purpose of: a) dispensing any economic poison; b) dispensing any other substance intended for plant nourishment, soil treatment, propagation of plant life, or pest control; or c) engaging in dispensing activities directly affecting agriculture, horticulture, or forest preservation, but not including the dispensing of live insects. (3) Economic poisons Economic poisons, as defined in part §137.3 of part 137, means: a) any substance or mixture of substances intended for preventing, destroying, repelling, or mitigating any insects, rodents, nematodes, fungi, weeds, and other forms of plant or animal life or viruses, except viruses on or in living man or other animals, which the Secretary of Agriculture declares to be a pest; or b) any substance or mixture of substances intended for use as a plant regulator, defoliant or desiccant. 201(d)(8) Operating limitations §36.1583(b) (1) Limitation on operations No person may operate, in accordance with the restrictions of §91.815 of part 91, an airplane that complies with the noise certification exemptions of this section, except: ### Page 55 a) to accomplish the work activity directly associ ated with the purpose for which it is designed; b) to provide flight crewmember training in the special purpose op eration for which the airplane is designed; or c) to conduct “nondispensing aerial work operations” related to agriculture, horticulture, or forest preservations. ### Page 56 Chapter 3 **TECHNICAL PROCEDURES APPLICABLE FOR NOISE CERTIFICATION OF ALL AIRCRAFT TYPES** **CHAPTER TABLE OF CONTENTS** 301 NOISE MEASUREMENT PROCEDURES … 58 301(a) Test Site Selection … 58 301(a)(1) Terrain … 58 301(a)(2) Grass … 59 301(a)(3) Snow … 59 301(a)(4) Plowed fields … 59 301(a)(5) Obstructions … 59 301(a)(6) Anomalous meteorological conditions … 60 301(b) Precautions Regarding Effects of Precipitation … 60 301(b)(1) General precautions … 60 301(b)(2) Effect of moisture on microphones … 60 301(b)(3) Microphone internal heaters … 60 301(c) Calibration of Acoustical Measurement System … 60 301(c)(1) General calibration process … 60 301(c)(2) Frequency-dependent system response corrections … 61 301(c)(3) Single-frequency level calibrations … 62 301(c)(4) Field sound pressure level calibration considerations … 62 301(c)(5) Application of calibration corrections … 63 301(c)(6) Calibration traceability … 63 301(d) Windscreen Insertion Loss … 64 301(e) Wind Speed Limitations … 64 301(f) Measurement of Atmospheric Conditions … 64 301(g) Meteorological Measurement System … 64 301(h) Aircraft Position Measurement … 65 301(i) Measurement System Specifications … 65 301(i)(1) Validity of measurent system configuration … 65 301(i)(2) Changes in measurement system configuration … 65 301(j) Noise Compliance Demonstration Plan … 65 302 FLIGHT PATH MEASUREMENT … 65 302(a) Radar or Microwave Tracking System … 65 302(a)(1) Aircraft equipment … 66 302(a)(2) Ground equipment … 67 302(b) Kine-theodolite System … 67 302(c) Radar/theodolite Triangulation … 67 302(d) Photographic Scaling … 67 302(e) DGPS-based Time-Space-Position Information Tracking Systems … 68 302(e)(1) General … 68 302(e)(2) System design issues … 69 302(e)(2)(A) Coordinate frames and waypoint navigation … 70 302(e)(2)(B) Test site survey … 70 302(e)(2)(C) Receiver output data … 71 302(e)(2)(C)(i) Data stored during aircraft noise testing when a real-time data link is used … 72 302(e)(2)(C)(ii) Data stored during aircraft noise testing when a real-time data link is not used … 72 302(e)(2)(C)(iii) Real-time DGPS messages … 72 302(e)(2)(C)(iv) Messages for multi-path testing … 73 302(e)(2)(D) System accuracy and sources of DGPS error … 73 302(e)(2)(E) Multi-path errors … 74 ### Page 57 302(e)(2)(E)(i) Characteristics … 74 302(e)(2)(E)(ii) Code-based system ground station … 74 302(e)(2)(E)(iii) Carrier-based system ground station … 75 302(e)(2)(E)(iv) Aircraft installation … 75 302(e)(2)(F) Other sources of DGPS error … 75 302(e)(2)(F)(i) Correction latency … 75 302(e)(2)(F)(ii) Tropospheric delay … 76 302(e)(2)(F)(iii) Mismatched GPS receivers … 76 302(e)(2)(F)(iv) Mismatched satellite ephemeris/clock data … 76 302(e)(3) System approval recommendations … 77 302(e)(3)(A) Design issues … 77 302(e)(3)(B) Data storage (logging) during noise testing … 78 302(e)(3)(B)(i) For a system with real-time data link … 78 302(e)(3)(B)(ii) For a system not using real-time data link … 78 302(e)(3)(C) Documentation … 79 302(e)(3)(D) Accuracy verification test … 79 302(e)(3)(E) Software verification … 79 302(e)(3)(F) Ground-station multi-path mitigation and verification … 79 302(e)(3)(F)(i) All systems … 80 302(e)(3)(F)(ii) Code-based systems … 80 302(e)(3)(G) Airport survey … 80 303 ON-BOARD FLIGHT DATA ACQUISITION … 80 303(a) General … 80 303(b) Magnetic Tape Recording … 81 303(c) Automatic Still Photographic Recording … 81 303(d) Cine Recording … 81 303(e) Video recording … 81 304 TIME SYNCHRONIZATION OF MEASURED DATA … 82 304(a) General … 82 304(b) TSPI Equipment and Software Approval … 82 304(c) Continuous Time-Dode Recording … 82 304(d) Recording of Single Time Marker … 83 304(e) Measurement of the Interval Between Recorder Start and Overhead … 83 304(f) Setting of Internal Time-Stamp Clock … 83 304(g) Additional Time-Synchronization Considerations … 83 305 CALCULATION OF CONFIDENCE INTERVALS … 84 305(a) Introduction … 84 305(b) Confidence Interval for the Mean of Flight Test Data … 84 305(b)(1) Confidence Interval for the Sample Estimate of the Mean of Clustered Measurements … 84 305(b)(2) Confidence interval for mean line obtained by regression … 85 305(c) Confidence Interval for Static-Test-Derived NPD Curves … 87 305(d) Confidence Interval for Analytically Derived NPD Curves … 88 305(e) Adequacy of the Model … 88 305(e)(1) Choice of engine-related parameter … 88 305(e)(2) Choice of regression model … 88 305(f) Worked Example of the Determination of 90 percent Confidence Intervals … 89 305(f)(1) Introduction … 89 305(f)(2) Confidence interval for a clustered data set … 90 305(f)(3) Confidence interval for a first order regression curve … 91 305(f)(4) Confidence interval for a second order regression curve … 95 305(f)(5) Confidence interval for the pooled data set … 96 305(g) Student’s t-distribution (for 90 percent Confidence) for Various Degrees of Freedom … 96 306 ADJUSTMENT OF AIRCRAFT NOISE LEVELS FOR THE EFFECTS … 97 306(a) Introduction … 97 ### Page 58 306(b) Definitions … 98 306(c) Background Noise Adjustment Procedure … 99 306(c)(1) Assumptions … 99 306(c)(2) Step-by step description … 100 306(c)(2)(A) Determination of pre-detection noise … 100 306(c)(2)(B) Determination of post-detection noise … 100 306(c)(2)(C) Testing of pre-detection noise versus post-detection noise … 100 306(c)(2)(D) Determination of masking criteria … 100 306(c)(2)(E) Identification of masked levels … 101 306(c)(2)(F) Determination of Last Good Band … 101 306(c)(2)(G) Adjustment of valid levels for background noise … 101 306(c)(2)(H) Adjustment of valid levels for measurement conditions … 101 306(c)(2)(I) Reconstruction of low frequency masked bands … 101 306(c)(2)(J) Reconstruction of levels for masked high frequency bands … 102 306(c)(2)(J)(i) Frequency extrapolation method … 102 306(c)(2)(J)(ii) Time extrapolation method … 103 306(c)(2)(K) Handling of spectra after reconstruction of masked bands … 103 306(d) General Considerations … 104 306(d)(1) Limitations and requirements for any background noise adjustment procedure … 104 306(d)(2) Rejection of spectra due to masking … 104 305(d)(3) Special tone correction considerations due to masking … 104 306(d)(4) Handling of masked data in reference conditions data sets … 105 307 NOISE REDUCTION SYSTEMS … 105 307(a) Variable Noise Reduction Systems … 105 307(a)(1) Reference procedures … 106 307(a)(2) Test conditions and procedures … 106 307(a)(3) Adjustments to measured noise data … 106 307(a)(4) Guidance for specific VNRS … 107 307(b) Selectable Noise Reduction Systems … 107 308 CALCULATION OF THE SPEED OF SOUND … 107 309 REFERENCE TABLES USED IN THE MANUAL CALCULATION OF … 107 301 NOISE MEASUREMENT PROCEDURES [ETM 3.1] The applicant should submit to the FAA for its review an d approval technical procedur es for the measurement of aircraft noise certification levels that the applicant will use. The technical procedures described in the following sections of this chapter should be considered as genera lly appropriate for all aircraft types. See Chapter 4 for procedures specific to Appendices A and H. See Chapter 5 for procedures specific to Appendix G. See Chapter 6 for procedures specific to Appendix J. 301(a) Test Site Selection [ETM 3.1.1] For airplanes, when the flight path intercept equivalent test procedure is used, and for helicopters, it may not be necessary for the test site to be located at an airport. De tails of the proposed noise certification test-site locations should be submitted to the FAA for review and approval. Some test-site criteria that could support selection of a nonairport test site include level terrain, reduced air traf fic, reduced ambient noise, improved weather conditions (temperature, humidity and wind), improved microphone placement, availability of field surveys, improved locations for aircraft position monitoring and improved pilot sight and handling. 301(a)(1) Terrain [ETM 3.1.1.1] Page 58 ### Page 59 Uneven terrain having features such as mounds or furrows can result in reflections that could influence the measured sound levels. Vegetation can reduce the amount of sound that is reflected from the ground surface. In most cases this effect results in a reduced sound level, but under some circumstances the level may be higher. Testing over a smooth hard surface, such as a paved area, will generally result in a higher sound level. 301(a)(2) Grass [ETM 3.1.1.2] For noise measurement points under the flight path 25 ft (7.5 m) radius circles of mowed grass (not exceeding 3 in (8 cm) height) are acceptable. For noise meas urement points located to the side of the flight path, the grass may be mowed in a semicircle of 25 ft (7.5 m) radius facing the line of flight. 301(a)(3) Snow [ETM 3.1.1.3] Snow in the area surrounding the noise measurement points may provide excessive absorption of aircraft sound reflected from the ground. Noise measurement points have been approved when snow within a 50 ft (15 m) radius of the noise measurement points has been removed. However, snow should not be piled at the borders facing the line of flight. 301(a)(4) Plowed fields [ETM 3.1.1.4] Earthen or sandy surfaces within a 25 ft (7.5 m) radius of the noise measurement points must be reasonably tamped down. Plowed furrows, silt or soft, powdered surfaces are unacceptable. 301(a)(5) Obstructions [ETM 3.1.1.5] Obstructions in the vicinity of the noise measurement points such as buildings, walls, trees, vehicles and test personnel, if close enough, may be unacceptable because of reflections that influence measured noise levels. There should be no obstructions that significantly influence the sound field from the aircraft with in a conical space above a point on the ground vertically below the microphone at each noise measur ement point. The cone is defined by an axis normal to the ground and by a half angle of 80º (75º for light propeller-driven airplanes) from the axis as illustrated in Table 2. Figure 2. Obstruction-free cone defined from the base of the measurements microphone Page 59 ### Page 60 301(a)(6) Anomalous meteorological conditions [ETM 3.1.1.6] Certain geographical areas are more susceptible to anomalous meteorological conditions than others (i.e., large variations, or inversions, of temperature or humidity, excessive turbulence, or thermally induced vertical winds). The applicant may conduct certification testing only as approved by the FAA. 301(b) Precautions Regarding Effects of Precipitation [ETM 3.1.2] 301(b)(1) General precautions [ETM 3.1.2.1] Fog, rain, drizzle and snow can have a number of adverse effects. Changes in sound generation and propagation under these conditions are not well documented. Most of the equipment used for measuring noise is not intended for use during conditions of precipitation, and the effects can range from changes in microphone and windscreen sensitivity or frequency response, to arcing of conventional condenser microphones, to possible failure of equipment because of electrical short circuits. 301(b)(2) Effect of moisture on microphones [ETM 3.1.2.2] Most microphones that are used during noise certification testing are susceptible to moisture. Precipitation, including snow, drizzle and fog, or excessive humidity may induce electrical arcing of the microphone sensors, making measured noise data unacceptable. Howeve r, some pre-polarized microphones are le ss susceptible to electrical arcing during high-moisture conditions (consult the equipment manufacturer’s specifications). Special care should be taken to ensure that any windscreens exposed to precipitation be thoroughly dry, inside and out, before use. Foam windscreens can trap water and wet foam windscreens should be avoided. 301(b)(3) Microphone internal heaters [ETM 3.1.2.3] When internal heaters are prov ided, microphones are less likely to be affected by moisture in wet, humid, cold or freezing atmospheric conditions. 301(c) Calibration of Acoustical Measurement System [ETM 3.1.3] 301(c)(1) General calibration process [ETM 3.1.3.1] The process for the calibration of an acoustical measurem ent system (see A36.3.1.1, A36.3.3.1, G36.103, and J36.109(c) of part 36 for definition of the “measurement system”) for purposes of noise certification testing consists of two parts: a) determination of system response corrections; and b) sound pressure level calibrations and testing in the field. Some of these calibrations must be conducted at the aircraft noise test site, with the measurement system configured and deployed for aircraft noise measurement. Note that additional calibrations and checks may be allowed, or required, by the FAA. ### Page 61 301(c)(2) Frequency-dependent system response corrections [ETM 3.1.3.2] Corrections for system response may include consideration of some or all of the following frequency-dependent elements to account for variations in sensitivity differences relative to the calibration check frequency (see A36.3.9.1 of part 36). Note.— To avoid confusion and to promote consistency, it is recommended that all response corrections be determined and presented in such a manner that the correction values are to be added to the measured aircraft noise levels to correctly adjust for the effects of system se nsitivity differences. Example: The correction value for a microphone response at a frequency wher e the microphone is more sensitive to sound than at the calibration check frequency should be presented as a negative number which, when added to the measured aircraft noise levels in that frequency band, lowers them to what would have been measured if the microphone response were flat. (1) Microphone free-field frequency response Corrections for microphone free-field frequency response may be determined using electrostatic actuator testing in combination with manufacturer-provided corrections for free-field conditions, or alternately from anechoic free-field testing using a method approved by the FAA (see A36.3.9.2 of part 36). (2) Microphone incidence corrections Microphone incidence corrections may be required when the microphone is not set at nominally grazing incidence. The corrections may be determined using manufacturer’s data, or alternately from anechoic free-field testing using a method approved by the FAA (see A36.3.9.3 of part 36). (3) Windscreen free-field insertion effects The windscreen free-field insertion effects may be determined using manufacturer’s data, or alternately from anechoic free-field testing using a method approved by the FAA (see A36.3.9.10 of part 36). Note.- If incidence corrections are required for microphone free-field response, then incidence corrections will likely also be required for the windscreen insertion effects. (4) System frequency response The frequency response corrections for the measurement sy stem as deployed in the field for aircraft noise measurements (not including the microphone and windscreen, but including any cables, attenuators, gain stages, signal conditioners, etc.) should be determined using a method approved by the certificating authority. Acceptable methods may include the use of discrete- or swept- si ne tones, or random or pseudo-random pink noise. The specific methods and techniques used should be submitted in advance to the FAA for approval. When using pink noise to determine system frequency response, there are additional specifications for noise generator performance and calibration specified in part 36. When using analog (direct or FM) magnetic tape recording, it is required that system frequency response be determined in the field while deployed for aircraft noise measurements (see A36.3.9.5 of part 36). (5) Effective bandwidth The effective bandwidth (also known as “bandwidth error” or “filter inte grated response”) for one-third octave band analyzer filters may be determined using manufacture r’s data, or from tests using discrete or swept sine signals as described in IEC 61260 (Reference T2). Note that bandwidth corrections might be redundant, dependent on the specific testing methodology used, if the measur ement system response, including the analyzer, is determined using pink noise or other broadband signal testing. ### Page 62 301(c)(3) Single-frequency level calibrations [ETM 3.1.3.3] The following calibrations are typically performed at the calibration check frequency: (1) Sound calibrator output level The output level of the sound calibrator should be dete rmined using means which are traceable to a national standards laboratory. This may include direct calibratio n by a certified metrology laboratory or comparison calibration methodology. Laboratory calibrations often include a certificate stating the sound pressure level, its tolerance, reference temperature and humidity, and the time period for which the calibration remains valid. This and other documentation may be required by the FAA for its approval. Sound calibrator output level corrections for barometric pressure, coupler volume, and other envi ronmental effects may be determined according to the manufacturer’s specifications. (2) Calibration of gain or attenuation settings The accuracy of gain or attenuation settings (range change s) for switchable components in signal conditioners, amplifiers, analyzers and sound level meters should be determined from manufactur er’s specifications or by laboratory testing, as approved by the FAA. (3) Sound pressure level sensitivity of the acoustical measurement system Sensitivity calibrations of the entire system must be performed before and after each day’s aircraft noise measurements using an approved sound calibrator. These calibrations should be performed in such a way that all system components, including cables, are accounted for. 301(c)(4) Field sound pressure level calibration considerations [ETM 3.1.3.4] It is necessary to establish and monitor the overall sound pressure level sensitivity of the measurement system while it is deployed in the field. Considerations for such field sound pressure level calibrations include: (1) Schedule for sound pressure level calibrations For aircraft noise certification tests defined in Appendices A, G, H and J, an acoustic calibration signal of known amplitude and frequency should be applied to the entire measurement system, including the microphone, while deployed in the field. This calibration should be performed as a minimum at the start and end of each day of aircraft noise measurements. When analog (direct or FM) magnetic recording media are used, the signal from the sound calibrator should also be recorded at the beginning of each physical volume of recording media (e.g. each tape reel, cartridge, or cassette) as well as at the end of the last physical volume of recording media. In addition, sound pressure level calibrations should be considered at regular intervals throughout each day of aircraft noise measurements (see A36.3.9.8, A36.3.9, and J36.109(e)(i) of part 36). As an example, sound pressure level calibrations are sometimes scheduled to coincide with meteorological flights between noise measurement runs, or during test aircraft refueling for longer test programs. Applicants should include the scheduling of such calibrations in the noise measurement test plan. (2) Aircraft noise data validity Aircraft noise data that are not preceded and succeede d by valid sound pressure level calibrations are not acceptable for aircraft noise certification purposes. Part 36 specifies that the difference between pre- and postcalibrations must not be greater than 0.5 dB. Because of this limitation, frequent checks of system sensitivity and functionality are advised (see A36.3.9.8 and J36.109(e)(2)(ii) Page 62 ### Page 63 (3) Insert “checks” Depending upon the specific measurement system config uration, and with the approval of the FAA, level sensitivity checks (e.g. charge or voltage “insert” calibrations, which involve electrical signals injected into the microphone preamplifier) at frequent intervals (e.g. every two hours of aircraft noise measurement deployment) may be used to supplement the requirement for sound pressure level calibrations at the start and end of each day of aircraft noise measurement. Employing such a procedure can facilitate continuous testing activity, and provide supporting data to isolate and identify potential system failures. Insert checks should not be considered as substitutes for required sound pressure level calibrations. (4) System configuration during sound pressure level calibrations in the field During field sound pressure level calibration, all components of the measurement system including cables, attenuators, gain and signal-conditioning amplifiers, filters (including pre-emphasis, if used) and power supplies, but excluding the windscreen, should be in place. Where used, attenuators and gain stages should be set to prevent overload and to maintain the calibration signal level within the reference level range specified in A3.6.6 of part 36. If any switchable filters that could affect the calibration signal are utilized during measurements, then sound pressure level calibrations should be performed both with and without these filters enabled. Components of the electrical system should not be added, removed or replaced without re-calibrating any part of the system affected by that component immediately before and after making each change. (5) Adjustments for differences between pre- and post- calibrations Section A3.9.8 of part 36 requires that the arithmetic m ean of the sound pressure level calibrations before and after a group of aircraft noise measurements must be us ed to represent the acoustical sensitivity level of the measurement system for those measurements. No such requirement is provided for in Appendices G or J of part 36. Nevertheless the FAA may approve such an adjustment to the system’s acoustical sensitivity level. In all cases, and as an alternative to the ar ithmetic mean of the before and after sound pressure level calibrations, the FAA may accept that for a particular aircraft noise ev ent the acoustical sensitivity level of the system be represented by a linear interpolation over time of the sound pressure level calibrations before and after the aircraft noise measurement to the time of the data acquisition for that aircraft event. 301(c)(5) Application of calibration corrections [ETM 3.1.3.5] It is recommended that all calibration corrections be determ ined and presented in a manner such that they are to be added to the measured aircraft levels to properly correct for system instrumentation effects (see Note to 301(c)(2)). All calibration corrections should be determined individually, reported with full documentation of the method of determination, and applied to the measured aircraft noise levels during data processing of analyzed one-third octave band sound pressure levels. This includes sound calibrator output adjustments for ambient conditions such as temperature and atmospheric pressure, coupler volume, etc. (see A36.3.9.7 of part 36), as well as any sensitivity “drift” corrections, or any system response corrections. Note 1.— Such adjustments should not be applied by adjusting the calibration value in the analyzer or sound level meter. Note 2.— In the event that the sound pressure level sensitiv ity of the analyzer or sound level meter needs to be reset after the initial calibration at the start of the test day, and after aircraft noise measurements have already been obtained, then prior to such reset, the sound pressure level calibration signal should be measured and recorded at the existing level sensitivity. In this way, a traceable record of the system sensitivity can be maintained. 301(c)(6) Calibration Traceability [ETM 3.1.3.6] Page 63 ### Page 64 All sound calibrator output levels and performance testin g of calibration instrumentation should be traceable to a national standards laboratory. The method of traceability may in some cases include documentation of any comparison calibration methods performed by the applicant, instrumentation manufacturer, or another third party. All methods of traceability must be approved by the FAA. 301(d) Windscreen Insertion Loss [ETM 3.1.4] The physical condition of a windscreen can significantly affect its performance due to insertion loss. Only new or clean, dry windscreens should be used. Insertion loss data adjustments for wi ndscreens used during aircraft noise evaluations under the provisions of Appendices A and H of part 36 may be obtained from manufacturer’s data or by free-field calibration in an anechoic chamber (see 401(c)(14) of this AC). For aircraft noise evaluations made under the provisions of Appendices G and J of part 36, windscreen insertion loss data adjustments may be required. 301(e) Wind Speed Limitations [ETM 3.1.5] The wind speed should be monitored against the specified wind speed limits. In cases where these limits are exceeded during an aircraft test run, that test run is invalid and might have to be repeated. No method has been approved for making data adjustments for wind speed or direction. 301(f) Measurement of Atmospheric Conditions [ETM 3.1.6] The temperature and relative humidity near the earth’s surface can be affect ed by numerous factors including solar heating, surface winds, local heating or cooling, increased or decreased local humidity, etc. To avoid localized anomalous conditions that often occur n ear the ground, meteorologi cal measurements are to be made 33 ft (10 m) above the surface for aircraft noise measurements made unde r the provisions of Appendices A and H of part 36 and between 4 ft (1.2 m) and 33 ft (10 m) for aircraft noise measurements made under the provisions of Appendices G and J of part 36. The meteorological conditions measured above the ground are assumed to be constant from the ground surface to the height at which they are measured. 301(g) Meteorological Measurement System [ETM 3.1.7] Atmospheric conditions should be measured within 6562 ft (2000 m) of the microphone locations and must be representative of the conditions existing over the geographical area in which noise measurements are conducted. If an applicant can show that measurements from a fixed meteorological station, such as might be found at a nearby airport comply with these requirements, then subject to the approval of the FAA, this facility may be used. Approval will normally require the applicant to show that the measurement systems have been calibrated within 90 days prior to the tests. In general, an approved portable system is preferred. This will be especially important when any of the meteorological conditions, in particular wind speed, are near their limits. Applicants should note that in order to determine the crosswind component an accurate measure of not only wind speed but also wind direction is required. Some wind direction indicators installed at airfields and airports have a slow response to rapid changes in wind direction and are not sufficiently accurate. ### Page 65 302 301(h) Aircraft Position Measurement [ETM 3.1.8] Appendices A, G, H and J specify that the aircraft position must be determined by a method independent of normal flight instrumentation. For this purpose the FAA has approved differential global positioning systems (DGPS), radar tracking, theodolite triangulation or photographic scaling techniques (see 302). 301(i) Measurement System Specifications [ETM 3.1.9] 301(i)(1) Validity of measu rent system configuration [ETM 3.1.9.1] Each applicant should submit information to the FAA about the measurement system used. The FAA will determine whether any listed components require approval. 301(i)(2) Changes in measu rement system configuration [ETM 3.1.9.2] If an applicant makes changes to the approved measurement system configuration, the FAA should be notified before aircraft noise certification testing to determine whether additional evaluation and approval are required. 301(j) Noise Compliance Demonstration Plan [ETM 3.1.10] An applicant should prepare a noise compliance demonstration plan (see 20 1(d)(1) of this AC), that specifies the proposed certification process, including the use of any equivalencies. This plan is to be submitted to the FAA allowing sufficient time to permit adequate review and possible revisions prior to the start of any noise certification testing. FLIGHT PATH MEASUREMENT [ETM 3.2] The criteria for the measurement of aircraft height and late ral position relative to the intended track are described in A36.2.3.1, F36.109(f)(6), G36.109(f), H36.101(d)(2), and J36.101(d)(2) of part 36. Examples of methods used include: a) radar tracking system; b) theodolite triangulation; c) photographic scaling; and d) differential global positioning system (DGPS)-based time-space-position information tracking systems. Practical examples of aircraft tracking systems employing one or more of these techniques are described in subsequent sections. Other tracking systems such as inertial navigation systems (INS) and microwave systems, which have a high degree of accuracy, have been installed in aircraft and consequently have been accepted by several FAA for use during noise certification. These techniques may be used singly or in combination. This is not intended to be an exhaustive list and additional information will be included as more experience is acquired. 302(a) Radar or Microwave Tracking System [ETM 3.2.1] One example of a radar position tracking system is shown in Figure 3. It operates on the principle of pulse radar with a Page 65 ### Page 66 radar interrogator (receiver/transmitter) located on the aircraft and a radar transponder (receiver/transmitter) positioned at each reference station. The elapsed time between the receiver/transmitter pulse and reception of the pulse returned from the reference station transponder is used as the basis for determining the range of each reference station. This range information, together with the known location of the reference stations, can be used to obtain a fix on the position of the aircraft in three dimensions. 4 32 1 Flight path Transponder site Microphone sites Acoustics test range position transponder layout K/T D BA C APATS positioning system with four reference stations Figure 3. Example radar/microwave position tracking system A pulse coding system is employed to minimize false returns caused by radar interference on reflected signals. The system performs the following basic functions during noise certification: a) continuously measures the distance between the aircraft and four fixed ground sites; b) correlates these ranges with In ter-Range Instrumentation Group Standard Serial Time Code Format B (IRIG B) time code and height information and out puts these data to a pulse code modulation (PCM) recorder; c) converts the aircraft range and height information into X, Y and Z position coordinates in real time; and d) uses the X, Y, Z data to drive a cockpit display providing the pilot with steering and position cueing. The accuracy of the coordinate calculation depends on the flight path and transponder geometry. Errors are minimized when ranges intersect, and the recommended practice is to keep the intersection angle near to 90. The four transponder arrangements shown in Figure 3 produce position uncertainties from ±3 ft to ±7 ft (±1.0 m to ±2.0 m). At low aircraft heights some inaccuracies can be introduced with the use of microwave systems. The use of a radioaltimeter can reduce these errors. The height data are recorded and synchronized with the microwave system. 302(a)(1) Aircraft equipment [ETM 3.2.1.1] The distance measuring unit computer and transponder beacon are connected to a hemispherical antenna which is mounted under the fuselage, on the aircraft centerline, preferably as close to the aircraft center of gravity as possible. ### Page 67 302(a)(2) Ground equipment [ETM 3.2.1.2] The four beacons should be located on either side of the aircraft ground track to permit an optimum layout. For example, a helicopter should be covered with angles between 30 and 150 (90 being the ideal angle). Two beacons can be located on the axis of the noise measurement points at distances of ±1640 ft (±500 m) from the central microphone, while another two beacons can be located on the track at ±1969 ft (±600 m) from the central microphone. 302(b) Kine-theodolite System [ETM 3.2.2] It is possible to obtain aircraft position data with classical kine-theodolites, but it is also possible to make use of a system composed of two simplified theodolites including a motorized photo-camera on a moving platform, which reports azimuth and elevation. These parameters are sync hronized with coded time and the identification number of every photograph recorded. Each 0.1-second azimuth and elevation data measurement is sent to a central computer which calculates the aircraft position (X, Y and Z) versus time for each trajectory. For example, for helicopter testing, photographic stations should be located at sideline positions about 984 ft (300 m) from the track, and at 656 ft (200 m) on either side of the three noise measurement points. The accuracy of such a system can be ±4.9 ft (±1.5 m) in X, Y and Z over the working area. 302(c) Radar/theodolite Triangulation [ETM 3.2.3] The opto-electronic system shown diagrammatically in Figure 4 uses a single optical theodolite to provide azimuth and elevation while range data are obtained from a radar tracking system using a single transponder. Data from these two sources are transferred to a desktop calculator at a rate of 20 samples/second from which three-dimensional position fixes can be derived. The system also provides tape start and stop times to the measuring sites, synchronizing all tape-recording times. The accuracy of the system is approx imately ±6.6 ft (±2.0 m), ±3.3 ft (±1.0 m) and ±6.6 ft (±2.0 m) for horizontal range (X), cross-track (Y) and he ight (Z) respectively. Uncert ainties associated with the determination of the visual glide slope indicator and ground speed are ±0.1 and ±0.5 kt (±0.9 km/h) 302(d) Photographic Scaling [ETM 3.2.4] The flight path of an aircraft during a noise certification demonstration may be determined by using a combination of ground-based cameras and height data supplied as a function of time from the on-board radio or pressure altimeters. For example, using this method for a helicopter, three camer as are placed along the intended track, such that one is sited close to the center microphone position and the othe r two are sited close to each of the 10 dB-down points, typically 1640 ft (500 m) either side of the microphone, depending upon the flight procedure being used. The cameras are mounted vertically and ar e calibrated so that the image size, obtained as the helicopter passes overhead, can be used to determine the height of the aircraft. It is important that the time at which each camera fires is synchronized with the on-board data acquisition system so that the height of the aircraft as it passes over each of the cameras can be correlated with the heights obtained from the photographs. The flight path of the aircraft as a function of distance may be obtained by fitting the aircraft data to the camera heights. ### Page 68 Generator 1.2 kW 240 V Interface Intercom Computer Dual floppydisc Radar transmitter/ receiver Optical tracking equipment Threshold detector Aircraft transponder Chart recorder Figure 4. Radar/optical position tracking system The aircraft reference dimension should be as large as possible in order to maximize photograph image size but should be chosen and used with care if errors in aircraft position are to be avoided. For a helicopter, foreshortening of the image due to factors such as main rotor coning (bending of the blades), disc tilt or fuselage pitch attitude, if not accounted for, will result in errors in the measurement of height and/or lateral and longitudinal position. By erecting a line above each of the camer as at right angles to the intended tr ack, at a sufficient height above the camera in order to provide a clear photographic image of both the line and the aircraft, the applicant may obtain the lateral offset of the aircraft as it passes over each of the cameras. This can be done by attaching marks to the line showing the angular distances from overhead at 5 intervals on either side of the vertical. For example, for helicopters, this method may be used to confirm that the helicopter follows a 6  ± 0.5 glide slope within 10 of the overhead of the center microphone as required by H36.107 of part 36. Furthermore, from the synchronized times of the aircraft passing over the three camera positions, the ground speed can be determined for later use in the duration adjustment. Overall accuracy of the system is ±1.0 percent of height and ±1.3 percent of longitudinal and lateral displacements. Mean approach/climb angles and mean ground speed can be determined within ±0.25 and ±0.7 percent respectively. 302(e) DGPS-based Time- Space-Position Information Tracking Systems [ETM 3.2.5] 302(e)(1) General [ETM 3.2.5.1] The use of conventional global positi oning system (GPS) receivers on-board aircraft to obtain time-space-position information (TSPI) is not considered to be accurate enough for noise certification testing. However, by using data Page 68 ### Page 69 from a second localized fixed-position GPS receiver, a substantial improvement in accuracy can be achieved. Such an arrangement is referred to as a differential GPS (DGPS) system. FAA may approve the use of such DGPS systems, based on the particular characteristics of the hardware, related software, installation and operational specifics proposed by the applicant. This section summarizes recommended requirements for DGPS systems proposed for use during noise certification testing. Typically the hardware components of these systems are GPS receivers and antennas on the ground and in the aircraft, the data link transmitter and antenna on the ground and the corresponding receiver and antenna in the aircraft, a laptop computer in the aircraft and batteries and electronic power supplies (see Figure 5). Software running on the laptop computer in the aircraft provides the user control/display function and performs data logging. A personal computer is generally needed to initialize the GPS receiver on the ground, but is not necessary for continuous operation. Personal computer Personal computer Applicant software Differential correction **CDI/GDI** A I R C R A F T G R O U N D GPS receiver GPS receiver Data link receiver Data link transmitter Figure 5. DGPS TSPI system basic architecture In addition to generating flight reference data for post-p rocessing, some applicants’ systems provide the pilot with information to navigate the aircraft. Measured aircraft position is compared to a desired reference flight path, and steering commands are sent to a course/g lide slope deviation indicator (CDI/GDI) installed specifically for use with the DGPS system. Variations on the basic architecture shown in Figure 5 are possible. For example, it is possible to eliminate the data link elements by collecting and storing data from both GPS receivers during a flight and post-processing these data in a single computer after the flight is complete. However, without a data link, DGPS data cannot be used for aircraft guidance, nor can an aircraft-based operator obtain “qui ck-look” information regarding the DGPS solution quality. Another possible variation on the basic architecture in Figure 5 involves the use of a two-way data link. Typically, identical transceivers would be used on the ground and in the aircraft. This enables gro und tracking of the aircraft during testing. 302(e)(2) System design issues [ETM 3.2.5.2] This section discusses DGPS system design issues including configuration, airport survey, DGPS receiver output data, and sources of error in DGPS systems. ### Page 70 302(e)(2)(A) Coordinate frames and waypoint navigation [ETM 3.2.5.2.1] The native coordinate system for GPS (i.e., the one in which its computations are performed) is the World Geodetic Survey of 1984 (WGS-84). Most GPS receivers provide output position information (latitude, longitude and altitude) in a variety of geodetic coordinate systems by transforming the WGS-84 position data. Aircraft noise certification tests typically involve the use of a rectangular coordinate frame whose definition is based upon the array of microphones or the centerline of an airport runway. Typically, the frame’s x-axis is established from two points on the ground that are nominally aligned with the runway centerline; the y-axis is orthogonal to the x-axis and also level, and the z-axis is vertical. Some GPS receivers can furnish data in a rectangular coordinate system based on waypoints. These are user-defined reference points intended to facilitate navigation along a route or in a local area. If a receiver supports waypoint navigatio n then two such points, defined in te rms of latitude, longitude and altitude, can be entered into the receiver. 1 The receiver will subsequently provide airc raft position relative to the coordinate frame implicitly defined by the points (i.e., the distance from the line connecting the two points and the distance to one point). If waypoint navigation is to be used for noise testing, then the initial survey performed to determine the position of the two waypoints is critical to the accuracy of the TSPI results (see 302(e)(2)(b)). If waypoint navigation is not available, or is not to be used, then the geodetic position solution (i.e., latitude, longitude and altitude) must be transformed to a local coordinate system through post-processing by the applicant prior to noise data processing. 302(e)(2)(B) Test site survey [ETM 3.2.5.2.2] A careful survey of the airport and near by areas where noise testing is to be conducted is critical to the success of a measurement program. The following steps are involved in a survey: a) An initial reference location, including numerical va lues for its latitude, longitude and altitude, is selected and its coordinates are stored in a permanent file for record keeping. Normally the initial reference location will be a surveyed monument on th e airport upon which latitude and longitude are stamped. Often the monument will have been derived from a third-order survey, in which case geodetic position errors of the order of hundreds of meters are not uncommon. However, such errors have virtually no effect on the measurement of positions relative to that point or another point derived from it. The published airport reference altitude can be assigned to the monument. Although this altitude typically is applicable to the base of the tower, the altitude difference between the mo nument and tower will not degrade the accuracy of differential measurements relative to the reference lo cation. Many GPS receivers have a “survey” mode whereby they average position measurements over a user-selected period of time (e.g., 24 hours) to generate a surveyed position estimate. Typical resulting absolute accuracies are 3 to 10 ft (0.9 to 3 m), which are more than adequate if the DGPS-based TSPI system meas urements will not be related to measurements from another system.2 1. For noise certification testing it is recommended that the GPS receiver read the waypoints from a printable data file. Alternatively, the waypoints could be keyed into the receiver and then written to a data file. 2. Prior to the advent of satellite-based techniques in the 1990s, land surveys were performed using an optical theodolite (to measure angles) and a chain (to measure linear distance). Networks of interlocking tria ngles were surveyed, with measur ements collected at each ver tex. The accuracy of such a survey was classified by the amount that the sum of the interior angles of a triangle deviated from 180º (af ter accounting for the earth’s curvature). A first-order su rvey was the most accurate; the vertices were typically 16 to 64 km (10 to 40 miles ) apart, and the angular error 1 arc second or less. Also, fo r a first-order survey, the latitude/longit ude of one point was measured by astrono mical means (accuracy approximately 15 m (50 ft)). A sec ond-order survey had vertices 8 to 16 km (5 to 10 miles) apart and maximum angular error of 5 arc minutes. A third-order survey had vertices 1.6 to 3.2 km (1 to 2 miles) apart and angular error not exceeding 15 arc minutes. ### Page 71 b) The DGPS-based TSPI system, with the ground-statio n antenna at the initial reference location, is used to measure the coordinates of the location where the ground-station antenna will be installed for the remainder of the test series. The latitude, longitude and altitude of this second location is stored in a permanent file for record keeping. If convenient, the ground-station antenna may be installed at the initial reference location for the duration of the test series. c) If waypoint navigation is to be used for the meas urement program, the DGPS-based TSPI system, with the ground station at the second (i.e., normal) location, is used to measure the latitude, longitude and altitude of the FROM and TO waypoints which will be used to esta blish the test program coordinate frame. At least three measurements should be made to guard against errors. The resulting locations should be stored in a permanent file for record keeping. d) The DGPS-based TSPI system, with the ground station at its normal location, is used to measure the microphone positions. The measured positions are stored in a permanent file for record keeping. If waypoint navigation is to be used for the measurement program then microphone positions should be recorded in test coordinates; otherwise latitude, longitude and altitude should be used. e) If it is not feasible to use the DGPS-based TSPI sy stem to survey the micropho ne locations, then direct measurements of at least three common points should be performed in order that the relationship between the two surveys can be determined. For example, if the microphones are surveyed us ing classical techniques, then a DGPS-based TSPI survey of the two microphones at the ends of a microphone line and one other microphone, as far removed from the first two as possi ble, will be sufficient. The surveys should agree to within 1 ft (30 cm) at each common point. If they differ by more than 1 ft (30 cm) and the difference can be expressed in terms of an offset and a rotation, then it may be possible to adjust the results of one survey to agree with the other. Such adjustments should be approved by the FAA prior to testing. The above tests should be performed, as a minimum, befo re and after each measurement program. Post-test data analysis should include a comparison of the two surveys. 302(e)(2)(C) Recei ver output data [ETM 3.2.5.2.3] This section addresses the GPS receiver messages3 (output data) which are of interest. All data are typically furnished via RS-232 serial ports (acceptable GPS receivers generally have multiple RS-232 ports). Three kinds of GPS receiver output data are of interest: a) data stored during flight testing for use during post-t est processing of noise data , collected from either the aircraft receiver when a real-time data link is used or from both receivers when a real-time data link is not used; b) differential correction data output by the ground-statio n receiver, transmitted to th e aircraft via a real-time data link, and input to the aircraft receiver. These data are not stored, but directly influence the accuracy of the stored data addressed in a); and c) data collected from the ground-station GPS receiver during multi-path verification tests prior to flight testing. 3. Standards organizations and manufacturers employ different terminology for pre-defined groups of data parameters available f rom receiver output ports. For example, in the United States, the National Marine Electronics Association (NMEA) uses the term “sentences,” the Radio Technical Commission for Maritime Services (RTCM) uses “message s,” Novatel Communications uses “logs” and Trimble Navigation Lt d. uses “cycle printouts.” ### Page 72 302(e)(2)(C)(i) Data stored du ring aircraft noise testing when a real-time data link is used [ETM 3.2.5.2.3.1] GPS receivers provide TSPI data in a variety of formats, both industry-standard and proprietary. In the United States, the National Marine Electronics Association (NMEA) has issued standards (References T7 and T8) which are intended to facilitate user communications with GPS receivers and other navigation de vices. Some GPS manufacturers have adopted NMEA standards, some use pr oprietary formats, and some use both . Manufacturers that provide NMEA outputs generally implement only a subset of the full set of messages set forth in the standards, and some follow the older Version 1.5 (Reference T7) rather than Version 2.0 (Reference T8), upon which this guidance was based. GPS receiver manufacturers have chosen different parameters to indicate the quality or status of the TSPI data. DGPS- based TSPI systems considered for noise certification tests, using a real-time data link, should save data from the aircraft GPS receiver in the receiver’s na tive raw format, in permanent files, for record keeping. Stored data should include time (e.g., Coordinated Universal Time (UTC) or GPS time with or without a local offset), aircraft latitude, longitude and altitude, or equivalently, aircraft position relative to a pre-defined waypoint, together with a status or quality flag indicating the reliability of the DGPS solution. Typically the applicant will employ post-processing software which will read the raw data, parse and format these data, perform any necessary transformations, and generate a file which will be used for noise data processing. Storage of raw data allows the FAA to verify the validity of the post-processed results. 302(e)(2)(C)(ii) Data st ored during aircraft noise testing when a real-time data link is not used [ETM 3.2.5.2.3.2] DGPS-based TSPI systems considered for noise certification tests which do not use a real-time data link should save data from both the ground and aircraft GPS receivers in raw (i.e., the receiver’s native) format, in permanent files, for record keeping. Manufacturers’ proprietary formats should be used since NMEA standard messages do not support this application. For post-processing, stored data should include time (e.g., UTC or GPS time) with or without a local offset, satellite ephemeris (see 302(e)(2)(F)(iv)) for a discussion of satellite ephemeris/clock data), pseudo-ranges,4 signal-to-noise ratios,5 and carrier phase.6 Applicants using dual-frequency (L1/L2) receivers will typically also save L2 carrier phase data.7 Typically, post-processing of the ground-based and airb orne GPS data will be performed using manufacturersupplied software. If this is not the case, then any applicant-developed software should be approved by the FAA. 302(e)(2)(C)(iii) Real-time DGPS messages [ETM 3.2.5.2.3.3] GPS receiver manufacturers have implemented both industry-standard and proprietary messages for use on real-time DGPS data links. The Radio Technical Commission for Maritime Services (RTCM), Special Committee 104 (SC - 104), has issued a standard (Reference T10) that is followed by most manufacturers. Manufacturers usually implement 4. Pseudo-range is the receiver’s measured distance to a satellite and is derived from the Coarse/Acquisition (C/A) code. It in cludes a receiver clock bias error and may be quantified in units of time or distance. 5. Signal-to-noise ratio (also called carrier-to-noise ratio) is derived from the receiver’s tracking loop circuits and is a me asure of the received signal strength. It is usually quantified in dB-Hz and varies from approximately 33 to 50. 6. Carrier phase is the amount of carrier cycles (at 1 575.42 MHz) which have accumulated since logging of this parameter was begun. It may be quantified in radians, degrees, cycles or feet (to convert to cycles, divide by the wavelength, 0.6247 ft). 7. The highest accuracy DGPS systems employ the signal carrier (L2 = 1 575.42 MHz), rather than the code (L1 = 1.023 MHz), which modulates the carrier, as the basic measurement observable. These technique s require that the number of full carrier cycles, i.e., 8-inch wavelengths, between the ground station and aircraft be determined once during a test. After the cycle count is established, the ground-station/aircraft separation is tracked to fractions of a wavelength, provided that the receiver carrier tracking loops (circuits) maintain phase lock. ### Page 73 only a subset of the RTCM/SC-104 messages, and some follow the older Version 2.0 of Reference T10 rather than Version 2.1, upon which this guidance was developed. Some manufacturers have also implemented proprietary DGPS messages which frequently bear a close resemblance to the RTCM/SC-104 messages. For applicants implementing a real-time DGPS data link, it is preferred that RTCM/SC-104 messages be employed for this purpose. Type-1 or Type-9 messages, each of which contains the actual DGPS corrections, should be selected and transmitted at a rate of 0.5 Hz or higher. Other message types (e.g., Type-3 ground-station location and Type-5 satellite health) may be used but should be sent at a rate of once per minute or slower. There is no recommended requirement for storing real-time DGPS correction data. The da ta status or quality flag (see 302(e)(2)(C)(i)) should however provide an indication that the correction data have been properly received and processed by the aircraft. 302(e)(2)(C)(iv) Messages for multi-path testing [ETM 3.2.5.2.3.4] Applicant-designed systems using code-based DGPS processi ng should collect and save data from dedicated multipath tests to be conducted prior to aircraft noise testin g (see 302(e)(2)(E)). Data co llected during multi-path tests should include individual satellite pseudo-ranges and signal-to -noise ratios. These parameters are provided only by receiver manufacturers’ proprietary messages. It is not necessary for applicants to conduct a dedicated test for systems using carrier-based DGPS processing. 302(e)(2)(D) System accuracy and sources of DGPS error [ETM 3.2.5.2.4] If only divergence (spherical spreading) of the noise is considered, and atmospheri c absorption mechanisms are ignored, then a 0.1 dB change in the noise level corresponds to a change of approximately 1.1 percent of the distance between the aircraft noise source and th e measurement microphone. Thus, for an aircraft altitude of 400 ft (122 m), the approximate minimum altitude during noise certification tests, a position error of 4.3 ft (1.3 m) along the line-ofsight vector connecting the microphone and aircraft can be expected to introduce a 0.1 dB error in the processed noise data. A position error of 10 ft (3.0 m) along the line-of-sight vector can be expected to introduce a 0.23 dB error in the processed noise data. For most DGPS systems, the most important error sources are, in decreasing order of importance, multi-path, correction latency and tropospheric delay. When these error sources are properly controlled, DGPS systems can provide accuracies between a few centimeters and approximately 15.1 ft (4.6 m) for an aircraft in low-dynamics flight regimes. Even the poorest of these accuracies is superior to that achieved by other conventional TSPI systems used for aircraft noise tests, including micr owave and photo-scaling. The best accuraci es are superior to those of a laser tracker. DGPS systems suitable for consideration by noise certification applicants can be expected to achieve an accuracy of a few inches to 5 ft (1.5 m). The highest accuracy is achieved using carrier-based techniques and post-flight processing of data collected from both the aircraft and ground-sta tion computers. Code-based solutions which use carrier smoothing (e.g., Novatel RT-20) achieve accuracies of 3.0 ft to 4.9 ft (0.9 m to 1.5 m), provided that the error sources discussed in this section are addressed properly. Consequently it is expected that the DGPS systems used for noise certification tests will introduce less than a 0.2 dB error into the noise data in a worst-case scenario (i.e., a noise certification approach measurement). Typical errors will be less than 0.1 dB for noise certification flyover and lateral measurements. In addition to the three error sources cited above, increases in DGPS position errors can also occur when the manufacturer and model of the ground station and aircraft GPS receivers are not the same, or when the ground station and aircraft receivers use different satellite ephemeris/clock data to specify the satellite orbital parameters. Sections 302(e)(2)(E)and 302(e)(2)(F) address all of the above errors and include methods for minimizing these errors or eliminating them entirely. ### Page 74 302(e)(2)(E) Multi-path errors [ETM 3.2.5.2.5] 302(e)(2)(E)(i) Characteristics [ETM 3.2.5.2.5.1] Multi-path refers to signals from GPS satellites which are reflected from objects (e.g., the ground, buildings and aircraft structural elements) before reaching the GPS antenna. Multi-path signals add algebraically to the desired lineof-sight signal and thereby decrease the accuracy of meas urements made with the latter. Multi-path conditions can occur independently at the aircraft and ground station antennas. Thus the differential correction data from the ground are not directly useful for correction for multi-path errors at the aircraft antenna. Rather, the broadcast corrections can contain ground-station multi-path errors which, in a statistical sense, add to those in the aircraft. Measurements have consistently shown that the presence of multi-path conditions at the ground station is significantly more deleterious than at the aircraft. This is because ground-station multi-path conditions vary slowly, acting like a bias over a test run of a few minutes, whereas the more dyna mic motion of the aircraft causes the effects of airborne multi-path conditions to behave like noise, which can be reduced somewhat by processing techniques such as filtering and averaging. For code-based processing, ground station multi-path error is typically between 1.0 ft and 10 ft (0.3 m and 3 m). Under very adverse conditions (e.g., GPS antenna near the side of and well below the top of a large building) multi-path errors can be several hundred meters. Multi-path errors associated with carrier-based processing techniques are significantly less than those for code-based methods and are usually of the order of centimeters. The extent of the multi-path error primarily depends on two factors: the capability of the ground station antenna and the location of the ground station antenna relative to reflec ting objects such as paved runways, buildings and parked aircraft. Receiver processing (e.g., the us e of narrow correlators available in most Novatel receivers) and/or carrier smoothing, available from several manufacturers, can reduce multi-path errors. 302(e)(2)(E)(ii) Code-based system ground station [ETM 3.2.5.2.5.2] To mitigate the effects of multi-path conditions on DGPS-based TSPI performance, the applicant’s ground station installation should meet the following requirements: a) the ground station should employ a multi-path-limiting antenna, such as one with a choke ring or an absorbing ground plane; and b) the ground station antenna should be mounted on a pole or tower, with unobstructed visibility of the sky. A minimum height of 10 ft (3 m) above ground level is recommended for the ground-station antenna. Additionally, to ensure that significant undetected multi-path errors do not corrupt the TSPI data collected during aircraft noise testing, the applicant’s ground-station inst allation should be tested for adequate multi-path condition performance prior to commencing the flight test. This can be done by collecting GPS receiver data during the same hours of the day that the system will be used for noise tests, with additional one-hour buffers on either side of this period. The data are then examined on a per-satellite basis, rather than navigation solution basis, for multi-path signatures. This examination should include at least pseudo-ranges and signal-to-noise ratios. If multiple periods of significant multi-path errors (i.e., several feet) are found, then a new location for the groundstation antenna should be selected and te sted. If only one or two isolated, brief multi-path incidents are found, then the antenna location can be retained but aircraft testing should not be conducted during these periods. Note.— The satellite-user geometry repeats over a cycle of approximately 23 hours 56 minutes. Thus if a ground Page 74 ### Page 75 station multi-path incident is observed one day, it is expected that a similar incident will occur 4 minutes earlier the following day. These procedures are similar to those utilized by the United States Coast Guard in checking out a marine DGPS station installation (References T10 and T19).8 After establishing a ground-station antenna site/configuration that satisfies the multi-path conditions criterion, the ground-station antenna should not be moved without performing another multipath test. The ground-station GPS receiver and any computer used in conjunction with the receiver may be removed and reinstalled without repeating the multi-path test. The multi-path verification test data should be saved as part of the permanent test-series data archive and should be made available for inspection by the FAA. 302(e)(2)(E)(iii) Carrier-bas ed system ground station [ETM 3.2.5.2.5.3] To mitigate the effects of multi-path conditions on DGPS-b ased TSPI performance, the applicant’s ground-station installation should meet the following recommended specifications: a) the ground station should employ a multi-path-limiting antenna, such as one with a choke ring or an absorbing ground plane; and b) the ground station antenna should be mounted on a pole or tower, with unobstructed visibility of the sky. A minimum height of (9.8 ft (3.0 m) above ground level is recommended for the ground-station antenna. There is no recommended requirement for collecting data to assess multi-path errors when carrier-based processing is employed. 302(e)(2)(E)(iv) Ai rcraft installation [ETM 3.2.5.2.5.4] It is expected that aircraft manufacturers will select a location on each aircraft model that minimizes multi-path effects. In this regard no recommended specifications have been developed. For most smaller aircraft (e.g., 10 seats or fewer), it has been found that the roof area directly behind the windshield is most advantageous. Manufacturers of larger aircraft have found forward positions on the roof to be desirable, although some have mounted the GPS antenna on the tail structure. Selecting a location for the GPS antenna on a helicopter may be more challenging since the main rotor will momentarily obscure most areas on the airframe. 302(e)(2)(F) Other sources of DGPS error [ETM 3.2.5.2.6] 302(e)(2)(F)(i) Correction latency [ETM 3.2.5.2.6.1] Correction latency, also called staleness, refers to the delay between the time of validity of a differential correction at the ground station and the time that the correction is applied in the aircraft. Delays in pro cessing at both ends of the ground-to-air data link can cause stale corrections to introduce unacceptably large errors. A second form of latency, solution latency, refers to the delay between the time at which a GPS receiver’s measurement is valid and the time when it is available at the output of the receiver. Solution delays are inherently smaller than correction 8. Coast Guard DGPS ground stations employ two GPS receiver/antenna pairs. The “additiona l” receiver/antenna pair (termed the i ntegrity monitor) provide a real-time continuous ch eck on the validity of the differential corrections generated by the “basic” receiver /antenna pair (termed the reference station). DGPS ground-station architectures be ing investigated for the FAA LAAS program employ between 2 and 4 receiver/antenna pairs to verify the corrections sent to the aircraft. No requirement for redundant ground-station equipment is recommended for DGPS-based TSPI systems used in noise certification tests. ### Page 76 delays and, in this context, are of concern only for aircraft guidance. For a system with a real-time data link which employs code-based DGPS solutions, it is strongly recommended that ground-to-aircraft messages conform to the RTCM/SC-104 standards used by the Coast Guard DGPS system.9 These messages contain pseudo-range rates-of-change, as well as th e correction at an identified time, to allow the user to correct for most of the latency-induced er ror. It is also preferred that the corr ections be computed and transmitted at least at a 0.5 Hz rate. 302(e)(2)(F)(ii) Tropospheric delay [ETM 3.2.5.2.6.2] The troposphere is that portion of the atmosphere between the earth’s surface and an altitude of approximately 20 miles (32 km). Differences in meteorological conditions between the ground station and the aircraft can cause dissimilar changes in the propagation times of signals from a satellite to these two locations. The effect is most pronounced for low-elevation-angle satellites. Since these changes are no t common to the two locations, they are not removed by differential corrections. Such tropospheric effects can contribute up to 66 ft (20 m) of ranging error on GPS signals, which can translate into as much as 33 ft to 39 ft (10 m to 12 m) of positioning error if not modeled and corrected. In differential mode this positioning error is typically less than 6.6 ft (2 m). Approximately 90 percent of these tropospheric propagation-related errors are due to the hydrostatic, or dry, component of tropospheric delay. Experiments performed for the FAA LAAS program have found tropospheric differences to introduce DGPS errors between 1 ft and 3 ft (0.3 m and 0.9 m) when the aircraft was at 3000 ft (914 m) altitude, but only a few centimeters when the receiver antennas were at the same altitude. To reduce the effects of tropos pheric errors on DGPS-based TSPI systems used in noise certification tests, it is recommended that use of these systems be limited to the aircraft being within a lateral distance of 20 nm (37 km) and a height of 5000 ft (1524 m) relative to the ground station. If desired, the hydrostatic component of the tropospheric delay can be effectively removed with the application of the tropospheric delay model (Reference T9) developed by the Radio Technical Commission for Aeronautics (RTCA) as per ICAO Annex 10 Standards and Recommended Practices, al ong with local meteorological measurements at the ground station. The relevant portion of this model is driven by local barometric pressure and satellite geometry (i.e., elevation angle). Reference T18 provides a functional overview of the RTCA model, as well as comparisons with other tropospheric propagation delay models. 302(e)(2)(F)(iii) Mismatched GPS receivers [ETM 3.2.5.2.6.3] Experiments have shown that DGPS errors are increased when the GPS recei vers at the ground station and in the aircraft are not “matched” in terms of manufacturer and model. With mismatched receivers, errors are increased moderately (e.g., 1.5 to 3 times) compared to those when the receivers are matched and when the satellites are operating normally. When a rare, soft satellite-failure, or signa l degradation, occurs, errors of several thousand feet have been observed.10 It is required that applicants’ systems have the same manufacturer/model GPS receiver on the ground and in the aircraft. 302(e)(2)(F)(iv) Mismatched sa tellite ephemeris/clock data [ETM 3.2.5.2.6.4] 9. The United States Coast Guard DGPS system’s broadcast messages (as well as marine systems of other nations) include the rate-of-change of each pseudo-range error, in addition to the pseudo-range error at a reference time. The user’s receiver is required to apply an adjusted correction consisting of the broadcast pseudo-range error, plus its rate-of-change multiplied by the time elapsed between the time the adjusted correction is applied, and the validity time for the pseudo-range correction. 10. Beginning on or before 21 October 1993, some differential users with mismatched ground and aircraft receivers experienced position errors of thousands of feet. The GPS Joint Program O ffice (JPO) of the United States Department of Defense attr ibuted the cause to a “def iciency” in Coarse/Acquisition code broadcast by satellite SVN19. It announced that the problem was corrected on 10 January 1994. Official statements are found in Notice Advisory to NAVSTAR User (NANU) 343-93294, 396-93337 and 006-94010. ### Page 77 GPS satellite broadcasts include a navigation message in the form of 50 bits per second modulation superimposed on the pseudo-random codes used for ranging. Within the navigation message are data sets that describe the satellite orbit (i.e., ephemeris information) and clock. These data sets are transmitted every 30 seconds. The GPS Control Segment uploads multiple ephemeris and clock data sets to the satellites, typically once per day. Note.— The Control Segment is the ground-based portion of the total GPS system. It includes the Operational Control facility in Colorado Springs, Colorado, United States, where the satellite ephemeris and clock data are calculated, five worldwide sites which collect satellite broadcast signals and provide data to the operational control facility, and three locations from which new ephemeris/clock data are uploaded to the satellites. Satellites typically change their broadcast ephemeris and clock message every four hours. The ephemeris/clock data sets are used by a receiver to compute its own position and, in the case of a reference station, differential corrections for use by other receivers. For a DGPS system to achieve full accuracy, both the ground station and aircraft receiver must use the same ephemeris and clock data sets. Internal receiver logic ensures that the ephemeris and clock data sets used by a given receiver are consistent for each satellite. However, occasionally the ground and aircraft receivers may use different ephemeris/clock data sets unless measures are taken by the user to ensure that the sets match. Mismatched ephemeris/clock data sets can occur for several reasons, e.g., a receiver is too busy performing other tasks when the data sets change, or a receiver encounters an error while decoding new data and continues to use an old data set. The RTCM/SC-104 messages used by the Coast Guard DGPS system guard against mismatched ephemeris/clock data sets by including the issue of data (I OD), an eight-bit data set label broa dcast by each satellite, in the broadcast messages (References T10 and T19). User receivers which conform to the RTCM/SC-104 standards will not apply differential corrections unless the IOD from the satellite and the DGPS correction message agree. The applicant should ensure that the ground station and aircra ft use the same ephemeris and clock data sets during testing. One way is to purchase GPS receivers and select DG PS messages which cause this check to be performed automatically. Another way to ensure agreement between the ground and aircraft ephemeris/clock data sets is to store in a permanent file for record keeping, at a rate of once each 30 seconds, the IOD used by each receiver and compare the IODs during posttest processing. 302(e)(3) System approval recommendations [ETM 3.2.5.3] This section summarizes approval recommendations for DGPS-based TSPI systems proposed for use during noise certification tests. 302(e)(3)(A) Design issues [ETM 3.2.5.3.1] Each applicant’s TSPI system design should address the issues identified in Table 4. The applicant’s documentation (see 302(e)(3)(C)) should address each item in the table. ### Page 78 Table 4. TPSI systems design development issues Number Issue Major considerations 1 Selection of processing method (real-time versus post-test) Need for aircraft guidance, ability to check testrun quality 2 Selection of solution method (carrier versus code) Accuracy (favors carrier), robustness (favors code), cost (favors code) 3 Use of geodetic or waypoint coordinates Waypoints can simplify post-processing but are not available for all receivers 4 Selection of GPS receiver and antenna Items 1, 2, 3 and others (antenna multi-path control, data messages, solution latency, matched air/ground receivers, and IOD capability) 5 Selection of data link equipment (if real-time system) Assigned frequency, data rate, error detection/ correction, flexible interface 302(e)(3)(B) Data storage (logging) during noise testing [ETM 3.2.5.3.2] 302(e)(3)(B)(i) For a system with real-time data link [ETM 3.2.5.3.2.1] For applicants employing a real-time data link, the ground-station GPS receiver should output RTCM/SC-104 Type- 1 messages at a rate of 0.5 Hz or great er, which should be transmitted to and us ed by the aircraft GPS receiver. The applicant’s aircraft computer should collect data from the aircraft GPS receiver and generate permanent data files containing: a) the three-dimensional aircraft pos ition copied directly from the receiver’s data port (i.e., in raw/native form) and not processed; b) if waypoint navigation is used, the waypoints (i.e., latitude, longitude and altitude) used to define the local coordinate frame; c) the time (e.g., UTC or GPS time), with or without a local offset, associated with each sample of position data copied directly from the receiver’s data port; and d) the data quality/validity indication associated with each sample of position data. If waypoints are used they should be included in the header of each data file. New waypoints should not be able to overwrite existing waypoints. If new waypoints are defined then a new data file should be created. For consistency with the noise data collected during a certif ication test, it is recommended that the data associated with a), c) and d) above be saved in th e GPS receiver’s raw/native format at a rate greater or equal to 2 Hz, the rate associated with the noise data. However, if hardware limitations preclude following this recommendation, a sampling rate of 0.5 Hz or greater is acceptable. 302(e)(3)(B)(ii) For a system not using real-time data link [ETM 3.2.5.3.2.2] Page 78 ### Page 79 TSPI systems which do not use a real-time data link should save data from both the ground and aircraft GPS receivers in raw/native format in a permanent file for record keep ing. Manufacturers’ proprietary formats should be used. NMEA standard messages do not support this application. Stored data should include: time (e.g., UTC or GPS time) with or without a local offset, satellite ephemeris, pseudorange, signal-to-noise ratio, and carrier phase. If tropospheric delay is being modeled, as described in 302(e)(2)(f) then local meteorological conditions should be measured and stored as well. It is recommended that applicants using dualfrequency (L1/L2) receivers also save L2 carrier phase data. Typically, post-processing of the ground-based and airborne GPS data will be performed using manufacturer-supplied software. If this is not the case then any applicantdeveloped software should be approved by the FAA. 302(e)(3)(C) Documentation [ETM 3.2.5.3.3] The applicant should prepare and submit documentation which includes: a) System description. Identifies, at a minimum, the issues in Table 4. b) Hardware description. Model and version number of all system components, including DGPS receivers, antennas, transceivers and computer. c) Software description. Software functionality and capabilities, da ta file formats, hardware required and operating system. d) System setup and operation. Ground and aircraft installation of the system including antennas, operating procedures, site survey procedures, power requirements and system limitations. e) Validating of the installation. A method often used is to park the aircraft at a known surveyed location and to read its position from the DGPS system. From a comparison of the DGPS and surveyed positions the installation can be verified. This can be performed either at the test site or at another location, such as the aircraft home base. As a minimum this process should be performed at the start and end of each measurement program and preferably at the beginning and end of each measurement day. 302(e)(3)(D) Accuracy verification test [ETM 3.2.5.3.4] The applicant should perform a one-time verification of the system accuracy, based on a minimum of six aircraft flight-test runs which encompass the conditions (i.e., speed, altitude, range and maneuvers) for which the system will be later used as a reference. The accuracy verification te st should involve a comparis on of the DGPS-based TSPI system’s position data with those from an accepted referen ce, such as a laser tracker or another approved DGPS system. This test should be performed on the complete DGPS-based TSPI system developed by the applicant. It is not adequate for an applicant seeking system approval to simply cite prior approval of another applicant’s system designed around the same GPS receiver. 302(e)(3)(E) Softwa re verification [ETM 3.2.5.3.5] Prior to using the system during a noise measurement program, any applicant-developed software for data logging and post-processing used to obtain data listed herein should be approved by the FAA. The approved software should be placed under version management. 302(e)(3)(F) Ground-station multi-path mitigation and verification [ETM 3.2.5.3.6] Page 79 ### Page 80 303 302(e)(3)(F)(i) All systems [ETM 3.2.5.6.1] The ground-station GPS receiver antenna should have a ch oke ring, absorbing ground plane, or other multi-path - reducing technique. The antenna should be positioned on a pole or tower at a minimum height of 10 feet above ground level. 302(e)(3)(F)(ii) Code-based systems [ETM 3.2.5.3.6.2] Prior to each measurement program, applicants using code-based DGPS systems should perform a multi-path investigation using the ground-station receiver and antenn a, as described in 302(e)(2 )(E)(ii). The results of the investigation should be saved as part of the permanent te st-series data archive and be made available for inspection by the FAA. 302(e)(3)(G) Airport survey [ETM 3.2.5.3.7] Additional information on survey requirements may be found in 302(e)(2)(B). Prior to and after completion of each measurement program, the applicant should use the DGPS-based TSPI system to survey the locations of: a) if no other method of survey is used, all microphones and, if used, waypoints; or b) if another method of survey is used, a recommended minimum of at least three points common to both methods. Survey data should be stored as part of the measurement-program permanent archive. If two survey methods are used, the common points should be reconciled to an accuracy of 1 ft (0.3 m) and the adjustment procedure submitted to the FAA for approval. ON-BOARD FLIGHT DATA ACQUISITION [ETM 3.3] 303(a) General [ETM 3.3.1] It is necessary to obtain the values of a variety of flight and engine parameters during the noise measurement period in order to: a) determine the acceptability of noi se certification flight tests; b) obtain data to adjust noise data; and c) to synchronize flight, engine and noise data. Typical parameters would include airsp eed, climb angle, height/altitude, gro ss weight, flap position, landing gear position, jet-engine thrust (power) setting parameters (e.g., compressor rotor speed, engine pressure ratio and exhaust gas temperature), helicopter rotor speed, engine torque and propeller rotational speed. A number of methods for collecting this information have been employed: a) manual recording; b) magnetic tape recording; ### Page 81 c) digital recording; d) automatic still photographic recording; e) cine recording; and f) video recording. Clearly, when a large number of parameters have to be co llected at relatively short time intervals, it may not be practicable to manually record the data. Thus the use of one of the automatic systems listed in b) to f) becomes more appropriate. The choice of a particular system may be influenced by a number of factors such as the space available, cost and availability of equipment. For systems which optically record the flight deck instruments, care must be taken to avoid strong lighting contrast such as would be caused by sunlight, deep shadow and reflections from the glass fronts of instruments, which would make data unreadable. To avoid this, it may be necessary to provide additiona l lighting to “fill in ” the deep shadow regions. To prevent reflections from the front of instruments, it is recommended that light-colored equipment or clothing on the flight deck be avoided. Flight crews should be required to wear black or dark-colored clothing and gloves. Furthermore, for systems which record the readings of dials it is important that the recording device be as near as possible, directly in front of the instruments to avoid parallax errors. 303(b) Magnetic Tape Recording [ETM 3.3.2] Multi-channel instrumentation tape recorders designed fo r airborne environments are employed for continuous recording of flight and engine performance parameters. Typical recorders are compact intermediate/wide band and can take both one-half-inch and one-inch magnetic tapes with a 24 to 28 volt DC power requirement. Six tape speeds as well as direct and FM recording are available in a tape recorder weighing about 60 lb (27) kg. 303(c) Automatic Still Photographic Recording [ETM 3.3.3] Photographs of the flight deck instrument panel can be taken by using a hand-held 35 mm single-lens reflex (SLR) camera with an 85 mm lens and high-speed slide film. The indications on the instruments can be read by projecting the slides onto a screen. 303(d) Cine Recording [ETM 3.3.4] Cine cameras with a one frame per second exposure rate have been used to acquire flight deck data. Care must be taken in mounting the camera to ensure that all the instruments that have to be photographed are within the field of view. Typical film cassettes containing about 2000 frames have been used with a frame counter to allow film changes to be anticipated. 303(e) Video recording [ETM 3.3.5] Flight and engine performance parameters can be recorded with a video camera, although as with cine cameras, care must be taken to ensure that all the instruments that have to be photographed are within the field of view. The recorded information is played back using freeze-frame features to obtain individual instrument readings. ### Page 82 TIME SYNCHRONIZATION OF MEASURED DATA [ETM 3.4] 304(a) General [ETM 3.4.1] Sections A36.2.3.2 and H36.101(d)(3) of part 36 specifies that there be precise time synchronization between noise measurements and aircraft position. Several methods have been used, such as noting the synchronization time on a clock mounted on the instrument panel which itself is recorded by the data acquisition system. One such system uses a ground camera that operates a radio transmission which, when received by an aircraft, lights two high-intensity light emitting diodes (LEDs) mounted in an analog clock attached to the instrument panel. Other methods for acquiring and processing TSPI are described in subsequent sections. A common time base should be used to synchronize noise, aircraft tracking and meteorological measurements. TSPI data should be determined at half-second intervals throughout the sound-measuring period (i.e., within 10 dB of maximum tone corrected perceived noise level (PNLTM)) by an approved method that is independent from systems installed aboard and normally used to control the aircraft. During processing, measured TSPI data should be interpolated, over time, to the time of sound emission of each half-second acoustic data record within the 10 dB-down period. Although the simplified procedure requires adjustment of only the PNLTM record to the reference flight path, sound emission coordinates should be determined for each half-second record for use in background noise adjustment procedures and for determination of incidence-dependent free-field microphone and windscreen corrections. 304(b) TSPI Equipment and Software Approval [ETM 3.4.2] Some off-the-shelf TSPI equipment may require software enhancement to accommodate the specific installation. All TSPI equipment and software should be demonstrated to, and approved by, the FAA to ensure the system’s operational accuracy. 304(c) Continuous Time-Dode Recording [ETM 3.4.3] This method uses a time-code signal, such as IRIG B, which is a modulated audio-frequency signal used for encoding time-base data developed by the Inter-Range Instrumentation Group (IRIG). In this method, the time-code signals from individual generators that have been synchronized to a common time-base are continuously recorded by both the noise data recorder(s) and by the TSPI system during measurement test runs. Synchronization of multiple generators can be performed either physically, by interconnecting via cable, or by means of radio transmission. The transmitted continuous time-code signal can be recorded directly, or used either continuously or in bursts to maintain synchronization of an independent time-code generator which is being recorded directly. This method allows for high-quality continuous timecode recording when there are intermittent reception problems. Note 1.— Synchronization should be accomplished at the start of each measurement day and checked at the end of each measurement day to minimize the effects of generator time drift. Any such drift should be documented and accounted for in processing. Note 2.— GPS-based measurement systems are often used for acquisition of TSPI data. GPS receivers are capable of providing the user with precise time-base information broadcast from the GPS satellite system, in some cases eliminating the need for a separate time-keeping device in the TSPI system. Note 3.— For noise data recording or for non-GPS-based TSPI systems, dedicated IRIG B time-code generators are available that use the GPS signal to constantly update and maintain time synchronization. Use of such a universal broadcast time-base can greatly simplify the logistics of time synchronization between measurement systems. Note 4.— There are two available time-bases for GPS-based systems: GPS Time and Coordinated Universal Time Page 82 ### Page 83 (UTC), whose values differ by more than 10 seconds at any given instant. Although the GPS signal includes both timebases, not all GPS receivers give the user access to both. Therefore, the user should exercise caution in identifying which time-base is used by each instrument. Note 5.— Many acoustical data recorders provide separate annotation channels in addition to the normal data channels. These channels are often not suitable for recording a modulated time-code signal because of limitations on dynamic range or bandwidth. In such cases a normal data channel of the recorder should be dedicated to recording the time-code signal. Note 6.— When continuous time-code recording is used, a nalysis of the recorded acoustic data can be initiated by routing the time-code channel output into a time-code reader and triggering the analyzer based on readout time. 304(d) Recording of Single Time Marker [ETM 3.4.4] This method involves transmittal and recording of a radio “hack”, or tone, usually used to indicate the “recorders on” or “overhead” time instant. This method typically requires a dedicated channel on both the noise and the TSPI recording systems. When such a system is used, analysis can be triggered manually by an operator listening for the hack, or by a detector circuit responding to the tone. When the operator wishes to start analysis at a time other than that of the time marker, a stopwatch or delay circuit can be used to delay triggering of the analyzer. When manual triggering is employed, the operator sh ould use extreme care to perform the tr iggering as accurately as possible. Accuracy to within one-tenth of a second can be expected from a conscientious human operator. 304(e) Measurement of th e Interval Between Recorder Start and Overhead [ETM 3.4.5] This method of synchronization involves use of a stopwatch or elapsed-time indicator to measure the interval between start-up of the noise data recorder and the instant that the aircraft position is overhead the centerline noise measurement point. This method can be employed successfully as long as the operator exercises care in timing, the determination of the overhead instant is performed accurately, and the star t-up characteristics of the r ecorder (in both record and playback modes) are known and repeatab le. Some recorders have variable star t-up times that cannot be predicted. Such recorders are not suitable for this method of synchronization. 304(f) Setting of Internal Time-Stamp Clock [ETM 3.4.6] Many digital recorders maintain a continuous internal time-of-day function by encoding time data in the recorded data stream. This method uses a digital recorder’s sub-code time, synchronized to the time-base used for the TSPI data. As with the continuous time-code recording method, synchronization by this method should be checked at the beginning and end of each measurement day and any drift accounted for in processing. Unfortunately, the time-setting function on many recorders does not provide for the necessary precision. The “second” digits cannot be made to “tick” in synchrony with an external clock. Such recorders are unsuitable for this method of synchronization. 304(g) Additional Time-Synchronization Considerations [ETM 3.4.7] Regardless of the synchronization method used, all elements affecting time synchronization, such as analyzer start-up delay, head displacement between normal and annotation data channels on analog recorders, and delays in automated triggering circuits, should be identified, quantified and accounted for in analysis and processing. Whenever human response to a timing event is required, errors cannot be accurately predicted, and conscientious operation is required to minimize such errors. The use of automated methods is preferred. Other methods, or variants of the listed methods, may be appropriate, but the use of all methods and instrumentation is subject to prior approval by the FAA. ### Page 84 CALCULATION OF CONFIDENCE INTERVALS [ETM 3.5] 305(a) Introduction [ETM 3.5.1] The use of NPD curves requires that confidence intervals be determined by using a more general formulation than is used for a cluster of data points. For this more general case, confidence intervals may have to be calculated about a regression line for: a) flight test data; b) a combination of flight test and static test data; c) analytical results; or d) a combination thereof. Items b) and c) are of particular significance for noise certification of an aircraft model range and require special care when pooling the different sources of sampling variability. Sections 305(b) to 305(e) provide an insight into the theory of confidence interval evaluation. Application of this theory and some worked examples are provided in 305(f). A suggested bibliography is provided in Appendix 3 to this AC for those wishing to gain a greater understanding. 305(b) Confidence Int erval for the Mean of Flight Test Data [ETM 3.5.2] 305(b)(1) Confidence Interval fo r the Sample Estimate of the Mean of Clustered Measurements [ETM 3.5.2.1] If n measurements of EPNLs y1, y2,….,yn are obtained under approximately the same conditions and it can be assumed that they constitute a random sample from a normal population with true population mean, , and true standard deviation, σ, then the following statistics can be derived: ӯ = estimate of the mean s = estimate of the standard deviation of the mean = From these and the Student’s t-distribution, the confidence interval, CI, for the estimate of the mean, ӯ, can be determined as: ௦ ܫܥ ൌ ӯ േ ݐሺଵିഀ,ఋሻ √௡,, మ Where ݐሺଵିഀ,ఋሻ denotes the ሺ1 െ ఈଶሻ percentile of the single-sided Student’s t-test with ߜ degrees of freedom (for a మ clustered data set ߜ ൌ ݊ െ 1 and where ߙis defined such that 100ሺ1 െ ߙሻ percent is the desired confidence level for the confidence interval. In other words it denotes the probability with which the interval will contain the unknown mean, . For noise certification purposes, 90 percent confidence intervals are generally desired and thus f ݐଽହ,ఋ is used (see Table 5 for a listing of values of ݐଽହ,ఋ for different values of ߜ .) ### Page 85 305(b)(2) Confidence interval for mean line obtained by regression [ETM 3.5.2.2] If n measurements of EPNL (y1, y2,….,yn ) are obtained under significantly varying values of engine-related parameter (x1, x2,….,xn ) respectively, then a polynomial can be fitted to the data by the method of least squares. For determining the mean EPNL, , the following polynomial regression model is assumed to apply: ߤൌܤ ଴ ൅ܤ ଵݔ൅ܤ ଶݔଶ ൅⋯൅ܤ ௞ݔ௞ . The estimate of the mean line through the data of the EPNL is given by: ݕൌܾ ଴ ൅ܾ ଵݔ൅ܾ ଶݔଶ ൅⋯൅ܾ ௞ݔ௞ . Each regression coefficient (Bi) is estimated by bi from the sample data using the method of least squares in a process summarized as follows. Each observation (xi, yi) satisfies the equations: ݕ௜ ൌܤ ଴ ൅ܤ ଵݔ௜ ൅ܤ ଶݔ௜ଶ ൅⋯.൅ ܤ ௞ݔ௜௞ ൅ߝ ௜ ൌܾ ଴ ൅ܾ ଵݔ௜ ൅ܾ ଶݔ௜ଶ ൅⋯.൅ ܾ ௞ݔ௜௞ ൅݁ ௜, Where ߝ௜ and ݁௜ are, respectively, the random error and residual associated with the EPNL. The random error ߝ௜ is assumed to be a random sample from a normal population with mean zero and standard deviation . The residual (݁௜) is the difference between the measured value and the estima te of the value using the estimates of the regression coefficients and xi . Its root mean square value (s) is the sample estimate for . These equations are often referred to as the normal equations. ### Page 86 Table 5. Student's t-distribut ion (for 90 percent confidence) for various degrees of freedom Degrees of freedom () t.95, 1 6.314 2 2.920 3 2.353 4 2.132 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 >60 1.645 The n data points of measurements (xi, yi) are processed as follows: Each elemental vector ሺݔ௜ሻ and its transpose ሺݔ௜ᇱሻ are formed such that: ݔ௜ ൌሺ 1ݔ ௜ ݔ௜ଶ..ݔ ௜௞ሻ, a row vector; and ۇݔ 1 ଶ ௜ ۊݔ௜ݔᇱ ൌ ۈ ۋ ,a column vector. ௜ ۈ. ۋ ۈ. ۋ . ݔۉ௜௞ی A matrix ܺ is formed from all the elemental vectors xi for i = 1, ….., n. ܺᇱ is the transpose of ܺ .A matrix ܣ is defined such that ܣ ൌܺ ᇱܺ and a matrix ܣିଵ is the inverse of ܣ In addition, ݕ ൌሺ ݕଵݕଶ …ݕ௡ሻ, and ܾ ൌሺ ܾ଴ܾଵ …ܾ ௞ሻ, with ܾ determined as the solution of the normal equations: ݕ ൌ ܾܺ and ݕ′ܺ ൌܾܺ′ܺ ൌܾܣ , Page 86 ### Page 87 to give ܾ ൌܣ ିଵݕ′ܺ . The 90 percent confidence interval CI 90 for the mean value of the EPNL estima ted with the associated value of the engine-related parameter x0 is then defined as: ܫܥଽ଴ ൌ ݕሺݔ଴ሻേݐ ଽହ,ఋݒݏሺݔ଴ሻ, where ݒሺݔ଴ሻ ൌ √ݔ଴ܣିଵݔ଴ᇱ . ,Thusܫܥଽ଴ ൌ ݕሺݔ଴ሻേ.ݐ ଽହ,ఋݏ√ݔ଴ܣିଵݔ଴ᇱ where: — ݔ଴ ൌሺ 1 ݔ଴ݔ଴ଶ …ݔ ଴௞ሻ; — ݔ଴ᇱ is the transpose of ݔ଴ ; — ݕሺݔ଴ሻis the estimate of the mean value of the EPNL at the associated value of the engine-related parameter; — .ݐଽହ,఍ is obtained for ζ degrees of freedom. For the general case of a multiple regression analysis involving K independent variables (i.e., K + 1 coefficients) ζ is defined as ζ = n-K-1 (for the specific case of a polynomial regression analysis, for which k is the order of curve fit, there are k variables independent of the dependent variable, and so ζ = n-K-1); and ሺ௬೔ି௬ሺ௫೔ሻሻమ — ݏൌ √∑೔స೔సభ೙ ௡ି௄ିଵ , the estimate of ߪ the true standard deviation. 305(c) Confidence Interval for Static-Test-Derived NPD Curves [ETM 3.5.3] When static test data are used in family certifications, NPD curves are formed by the linear combination of baseline flight regressions, baseline projected static regressions and derivative projected static regressions in the form: EPNLDF = EPNLBF – EPNLBS + EPNLDS, or using the notation adopted above: ݕ஽ிሺݔ଴ሻ ൌ ݕ஻ிሺݔ଴ሻ െ ݕ஻ௌሺݔ଴ሻ ൅ ݕ஽ௌሺݔ଴ሻ, where: DF denotes derivative flight; BF denotes baseline flight; BS denotes baseline static; and DS denotes derivative static. ### Page 88 Confidence intervals for the derivative flight NPD curves are obtained by pooling the three data sets, each with their own polynomial regression. The confidence interval for the mean derived EPNL at engine-related parameter x0, i.e., for ߤ஽ிሺݔ଴ሻ, is given by: ܫܥଽ଴ሺݔ଴ሻ ൌ ݕ஽ிሺݔ଴ሻേ ݒ′ݐ ஽ிሺݔ଴ሻ where: ݒ஽ிሺݔ଴ሻ ൌ√ ሺ ݏ஻ிݒ஻ிሺݔ଴ሻሻଶ ൅ሺ ݏ஻ௌݒ஻ௌሺݔ଴ሻሻଶ ൅ሺ ݏ஽ௌݒ஽ௌሺݔ଴ሻሻଶ with sBF, sBS, sDS, vBF (x0), vBS (x0), vDS (x0) computed as explained in 3.5.2.2 for the respective data sets indicated by the subscripts BF, BS and DS, and ݐᇱ ൌ ሺݏ஻ிݒ஻ிሺݔ଴ሻሻଶݐ஻ி ൅ሺ ݏ஻ௌݒ஻ௌሺݔ଴ሻሻଶݐ஻ௌ ൅ሺ ݏ஽ௌݒ஽ௌሺݔ଴ሻሻଶݐ஽ௌ ሺݏ஻ிݒ஻ிሺݔ଴ሻሻଶ ൅ሺ ݏ஻ௌݒ஻ௌሺݔ଴ሻሻଶ ൅ሺ ݏ஽ௌݒ஽ௌሺݔ଴ሻሻଶ where tBF, tBS, tDS are the t95, ζ values each evaluated with the respective degrees of freedom ζBF, ζBS, ζDS as they arise in the corresponding regressions. 305(d) Confidence Interval for Analytically Derived NPD Curves [ETM 3.5.4] Analysis may be used to determine the effect of changes in noise source components on certificated levels. This is accomplished by analytically determining the effect of hardware change on the noise component it generates. The resultant delta (Δ) is applied to the original configuration, and new noise levels are computed. The changes may occur on the baseline configuration or on subsequent derivative configurations. The confidence intervals for this case are computed using the appropriate method from 305(b) and 305(c). If Δ̂ represents the analytically determined change and if it is assumed that it may deviate from the true unknown ∆by some random amount, d, such that: ߂̂ ൌ∆൅݀ , Where d is assumed to be normally distributed with mean zero and known variance ߬ଶ, then the confidence interval for ߤሺݔ଴ሻ ൅∆ is given by: ሺݕሺݔ଴ሻ ൅߂ ̂ ᇱݒሻേݐ ᇱሺݔ଴ሻ, Where ݒᇱሺݔ଴ሻ ൌ√ ݒ ሺ ݔ଴ሻଶ ൅߬ ଶ and t’ is as above without change. 305(e) Adequacy of the Model [ETM 3.5.5] 305(e)(1) Choice of engi ne-related parameter [ETM 3.5.5.1] Every effort should be made to determine th e most appropriate engine-related parameter, x, which may be a combination of various simpler parameters. 305(e)(2) Choice of regression model [ETM 3.5.5.2] Page 88 ### Page 89 It is not recommended in any case that polynomials of gr eater complexity than a simple quadratic be used for certification purposes, unless there is a clear basis for using a higher order polynomial. Standard texts on multiple regression should be consulted, and the data available should be examined to show the adequacy of the model chosen. 305(f) Worked Example of the Determination of 90 percent Confidence Intervals from the Pooling of Three Data Sets [ETM 3.5.6] 305(f)(1) Introduction [ETM 3.5.6.1] This section presents an example of the derivation of the 90 percent confidence intervals arising from the pooling of three data sets. Worked examples and guidance material are presented for the calculation of confidence intervals for a clustered data set as well as for first order (i.e., straight line) and second order (i.e., quadratic) regression curves. In addition this section also shows how the confidence interval must be established for the pooling together of several data sets. Consider the theoretical evaluation of the certification noise levels for an aircraft retrof itted with silenced engines. The approach noise level for the “flight datum” aircraft was derived from a clustered data set of noise levels measured at nominally reference conditions, to which were added source noise corrections derived from a quadratic least-squares curve fit through a series of data points made at different e ngine thrusts. In order to evaluate the noise levels for the aircraft fitted with acoustically treated engines, a further source noise curve, assumed to be a straight least-squares regression line, was established from a series of measurements of the silenced aircraft. Each of the three databases is assumed to be made up of data unique to each base. The clustered data set consists of six EPNL levels for the nominal datum hard wall condition. These levels have been derived from measurements which have been fully corrected to the hard wall approach reference condition. The two curves which determine the ac oustical changes are the regression cu rves (in this example quadratic and straight-line least-squares fit curves) for the plots of EPNL against normalized thrust for the hard wall and silenced conditions. These are presented in Figure 6 where the dotted lines plotted about each line represent the boundaries of 90 percent confidence. Each of the two curves is made up of the full set of data points obtained for each condition during a series of back-toback tests. The least-squares fits therefore have associated with them all the uncertainties contained within each data set. It is maintained that the number of data points in each of the three sets is large enough to constitute a statistical sample. ### Page 90 Corrected net thrust per engine 90.0 91.0 92.0 93.0 94.0 95.0 96.0 97.0 98.0 99.0 100.0 1 000 1 200 1 400 1 600 1 800 2 000 2 200 Hardwall Silenced ( = 89.93 + 0.001843 )y x EPNL Figure 6. Regression curves for plots of EPNL against normalized thrust for hard wall and silenced conditions 305(f)(2) Confidence interval for a clustered data set [ETM 3.5.6.2] The confidence interval for the clustered data set is defined as follows: Let EPNLi be the individual values of EPNL, n = number of data points; and t = Student’s t-distribution for ( n-1) degrees of freedom (i.e., the number of degrees of freedom associated with a clustered data set). Then the confidence interval ܫܥ ൌ ܮܰܲܧ േݐ √ ௦ ௡ , where s, the estimate of the standard deviation, is defined as: ሺா௉ே௅೔ିா௉ே௅ሻమݏൌ √∑೔స೔సభ೙ , and௡ିଵ ೔సభ ா௉ே௅೔ܮܰܲܧ ൌ ∑೔స೙ ௡ . ### Page 91 Suppose that the clustered set of EPNL values consists of the following: Run number EPNL 1 95.8 2 94.8 3 95.7 4 95.1 5 95.6 6 95.3 Then the number of data points (n) = 6, the degrees of freedom (n-1) = 5, and the Student’s t-distribution for 5 degrees of freedom = 2.015 (see Table 5 ), and so: ೔స೙ ா௉ே௅ ܮܰܲܧ ൌ ∑೔సభ ௡ ೔ ൌ9 5 . 3 8, ሺா௉ே௅೔ିா௉ே௅ሻమݏൌ √∑೔స೔సభ೙ ൌ 0.3869,௡ିଵ and the confidence interval (CI) is calculated as follows: ܫܥ ൌ ܮܰܲܧ േݐ √ ௦ ௡ ൌ 95.38 േ 2.015 ଴.ଷ଼଺ଽ ൌ 95.38 േ 0.3183√଺ 305(f)(3) Confidence in terval for a first order regression curve [ETM 3.5.6.3] Suppose that the regression curve for one of the source noise data sets, the silenced case, can best be represented by a straight line least-squares fit curve (i.e., a first order polynomial). The equation for this regression line is of the general form: Y = a + bX where Y represents the dependent variable EPNL, and X represents the independent variable, in this case normalized thrust ிಿ ఋ . Although for higher order polynomial least-squares curves a re gression line’s coefficients (i.e., the solutions to the “normal equations”) are best established through computer matrix solutions, the two coefficients for a straight-line fit, a and b, can be determined from the following two simple formulae for the measured values of X and Y, Xi and ∆௜ : ೔స೙ ೔స೙ ܽൌ ∑೔సభ ௒೔ି௕ ∑೔సభ ௑೔; and௡ మ ܾൌ ஼௢௩௔௥௜௔௡௖௘ మ , where:௏௔௥௜௔௡௖௘ ൌ ௌೣ೤ ௌೣ ೔స೙ ௑ ௒ ೔స೙ ௑ ∑೔స೙ ௒ ݕݔܵଶ ൌ ∑೔సభ ௡ ೔ ೔ െ ∑೔సభ ௡ ೔ మ ೔సభ ೔ ;and Page 91 ### Page 92 ೔స೙ మ ೔ష೙ ܵ௫ଶ ൌ ∑೔సభ ௑೔ ೔సభ ௑೔ሻଶ௡ െሺ ∑ ௡ . The 90 percent confidence interval about this regression line for X = x0 is then defined by: ܫܥଽ଴ ൌ ܻ േݏݐ √ ݔ଴ܣିଵݔ଴ᇱ where: t = Student’s t-distribution for 90 percent confidence corresponding to (n – k - 1) degrees of freedom, where k is the order of the polynomial regression line and n is the number of data points; ݔ଴ ൌሺ 1 ݔ଴ሻ; ᇱݔ଴ ൌሺ ݔ1଴ሻ; ܣିଵ is the inverse of ܣ where ܣ ൌܺ′ܺ , with ܺ and ′ܺ defined as in 305(b)(2) from the elemental vectors formed from the measured values of the independent variable ܺ௜; and ݏൌ√ ݅ൌ1݊െ݇െ1 Where ሺ∆ܻሻ௜ ൌ the difference between the measured value of Y i at its associated value of X i, and the value of Y derived from the straight line least-squares fit curve for X = Xi, and n and k are defined as the number of data points and the order of the polynomial regression line, respectively. Suppose that the data set consists of the following set of six EPNL values, together with their associated values of engine-related parameter (see Table 6). Note that it would be usual to have more than six data points making up a source noise curve, but in order to limit the size of the matr ices in this example, the number of data points has been restricted. By plotting this data (see Figure 6) it can be seen by examination that a linear relationship between EPNL (the dependent variable Y) and ி ఋ ಿ (the independent variable X) is suggested with the following general form: Y = a + bX. ### Page 93 The coefficients a and b of the linear equation are defined as above and may be calculated as follows: X 1 395 1 505 1 655 1 730 1 810 1 850 Y 92.3 92.9 93.2 92.9 93.4 93.2 XY 128 759 139 815 154 246 160 717 169 054 172 420 X2 1 946 025 2 265 025 2 739 025 2 992 900 3 276 100 3 422 500 ∑X ∑Y ∑XY ∑X2 9 945 557.9 925 010 16 641 575 ହହ଻.ଽିሺ଴ ೔స೙ ௒೔ି௕ ∑೔స೙ ௑೔ ଴଴ଵ଼ସଷሻሺଽଽସହሻܽൌ ∑೔సభ ௡ ೔సభ ൌ ଺ ൌ8 9 . 9 3and మ ܾൌ ஼௢௩௔௥௜௔௡௖௘ మ , where: ௏௔௥௜௔௡௖௘ ൌ ௌೣ೤ ௌೣ ೔స೙ ௑ ௒ ೔స೙ ௑ ∑೔స೙ ௒ ൌ ଽଶହ଴ଵ଴ െ ሺଽଽସହሻሺହହ଻.ଽሻ ݕݔܵଶ ൌ ∑೔సభ ೔ ೔ െ ∑೔సభ ௡ ೔ మ ೔సభ ೔ ൌ 48.46; and ௡ ଺ ଷ଺ ೔స೙ మ ೔ష೙ ೔సభ ௑೔ ೔సభ ௑೔ሻଶ ൌ ଵ଺଺ସଵହ଻ହ െሺ ଽଽସହ ܵ௫ଶ ൌ ∑ െሺ ∑ ଺ ሻଶ ൌ 26.289.6, to give: ௡ ௡ ଺ ܾൌ ସ଼.ସ଺ ଶ଺ଶ଼ଽ.଺ ൌ 0.001843. The 90 percent confidence interval about this regression line is defined as: ܫܥଽ଴ ൌ ܻ േݏݐ √ ݔ଴ܣିଵݔ଴ᇱ and is calculated as follows. From the single set of measured independent variables tabulated in Table 6, the matrix, ܺ ,is formed from the elemental row vectors such that: 1 1395 1 1505 ܺ ൌ ۇ 1 1655ۊ1ۈۋ ۈ 1730ۋ 1 1810 ۉ1 1850ی and ′ܺ ,the transpose of ܺ ,where 1 1 1 1 1′ܺ ൌሺ 11395 1505 1655 1730 1810 1850ሻ. The matrix ܣ is now formed, defined such that ܣ ൌ ܺ′ܺ and so: ܣ ൌሺ 6 99459945 16641575ሻ, and its inverse ܣିଵ such that: ܣିଵ ൌሺ 17.5836 െ0.01051െ0.01051 6.3396 ∗ 10ି଺ሻ. ### Page 94 Note.— The manipulation of matrices (i.e., their multiplication and inversion) is best performed by computers via standard routines. Such routines are possible using stan dard functions contained within many commonly used spreadsheets. Table 6. Values of sample data set Run number ܨ ߜ ே EPNL 1 1 395 92.3 2 1 505 92.9 3 1 655 93.2 4 1 730 92.9 5 1 810 93.4 6 1 850 93.2 To find the 90 percent confidence interval about the regression line for a value of ிಿ (i.e., x0) of 1 600, the row vector ሺݔ଴ሻ and its transpose ሺ′ݔ଴ሻ, a column vector, are formed such that: ఋ ݔ଴ ൌ ሺ1 1600ሻ, and ′ݔ଴ ൌሺ 11600ሻ. . From the calculation of ܣିଵ one obtains: ݔ଴ܣିଵ ൌ ሺ1 1600ሻሺ 17.5836 െ0.01051െ0.01051 6.3396 ∗ 10ି଺ሻ ൌ ሺ0.7709 െ 3.6453 ∗ 10ିସሻ, and so: ݔ଴ܣିଵ′ݔ଴ ൌ ሺ0.7709 െ 3.6453 ∗ 10ିସሻሺ 11600ሻ ൌ 0.1876. The equation for confidence interval also requires the value of standard deviation for the measured data set to be evaluated. From Table 6 and the regression equation for the least-squares best fit straight line, from which is calculated the predicted value of EPNL at each of the six measured values of ிಿ one proceeds as follows: ఋ Page 94 ### Page 95 Run number ܨ ߜ ே EPNL (Measured) 1 1 395 92.3 2 1 505 92.9 3 1 655 93.2 4 1 730 92.9 5 1 810 93.4 6 1 850 93.2 Run number EPNL (Predicted) ሺ∆ܮܰܲܧሻଶ 1 92.50 0.03979 2 92.70 0.03911 3 92.98 0.04896 4 93.12 0.04708 5 93.26 0.01838 6 93.34 0.01909 ଶ௜ୀ ௜ୀଵ ௡ሺ∆ܮܰܲܧሻ௜ ൌ√ 0.21241ݏൌ √∑ ݊െ݇െ1 6െ1െ1 ൌ 0.2304 for n = 6 and k = 1. Taking the value of Student’s t-distribution from Table 5 for ( n – k - 1) degrees of freedom (i.e., 4) to be 2.132, the confidence interval about the regression line at ி ఋ ಿ  = 1600 is defined as follows: ܫܥଽ଴ ൌ ܮܰܲܧ ൅ݏݐ √ ݔ଴ܣିଵ′ݔ଴ ൌ 92.88 േ ሺ2.132ሻሺ0.2304ሻ√0.1876 ൌ 92.88 േ 0.2128 In order to establish the lines of 90 percent confidence intervals about a regression line, the values of CI90 for a range of values of independent variable(s) should be calculated, through which a line can be drawn. These lines are shown as the dotted lines in Figure 6. 305(f)(4) Confidence int erval for a second order regression curve [ETM 3.5.6.4] The confidence intervals for a second order regression curve are derived in a similar manner to those for a straight line detailed in 305(f)(3). A detailed example of their calculation is not discussed here. However the following points should be borne in mind. The coefficients of the least-squares regression quadratic line are best dete rmined via computer matrix solutions. Regression analysis functions are a common feature of many proprietary software packages. The matrices ݔ଴, ′ݔ଴, ݀݊ܽ,ܺ ′ܺ formed during the computation of the confidence interval according to the formula: ### Page 96 ܫܥଽ଴ ൌ ܻ േݏݐ √ ݔ଴ܣିଵ′ݔ଴ are formed from 1 × 3 and 3 × 1 row and column vectors respectively, made up from the values of independent variable X according to the following general form: ᇱݔ ൌ ሺ1 ݔ ݔଶሻ݀݊ܽ ݔ ൌሺ ݔ 1 ݔ ଶ ሻ. The number of degrees of freedom associated with a multiple regression analysis involving K variables independent of the dependent variable (i.e., with ( K+1) coefficients, including the constant term) is defined as ( n – K - 1) . For a second order regression curve, there are two independent variables and so the number of degrees of freedom is (n-3) . 305(f)(5) Confidence interval for the pooled data set [ETM 3.5.6.5] The confidence interval associated with the pooling of three data sets is defined as follows: ܫܥ ൌ ܻ േܶ √∑௜ୀ௜ୀଵଷ ܼ௜ଶ , where: ൌ ஼ூ೔ܼ௜ ,௧೔ With CIi = confidence interval for the i-th data set, ti = the value of Student’s t-distribution for the i-th data set, and ೔సయ ܶൌ ∑೔సభ ௓೔మ௧೔ .∑೔సయ ௓మ೔సభ ೔ The different stages in the calculation of the confidence interval at the reference thrust of ி ఋ ಿ ൌ 1600 for the pooling of the three data sets are summarized in Table 7. 305(g) Student’s t-distribution (for 90 percent Confidence) for Various Degrees of Freedom [ETM 3.5.7] The values in the Student’s t-distribution to give a probability of 0.95 that the population mean value (µ) is such that: ߤݏ൑ ݕ ൅.ݐ ଽହ,ఋ √݊ and thus a probability of 90 percent that ݕ െ.ݐ ଽହ,ఋ √ ݏ ݊ ൑ߤ൑ݕ ൅.ݐ ଽହ,ఋ ݏ √݊ are tabulated in Table 5. ### Page 97 306 Table 7. Example of confidence interval calculation Description Reference thrust Function ܨ ߜ ே Datum Hard wall 1 600 Silenced 1 600 90% confidence interval about the mean CI90 0.3183 0.4817 0.2128 Number of data points n 6 23 6 Degree of curve fit k 0 2 1 Number of independent variables K 0 2 1 Number of degrees of freedom N – K - 1 5 20 4 Student’s t-distribution Z Z2 Z2t t ܫܥଽ଴ ݐ ሺܫܥଽ଴ ݐ ሻଶ ሺܫܥଽ଴ ݐ ሻଶݐ 2.015 0.1580 2.4953 × 10–2 5.0280 × 10–2 1.725 0.2792 7.7979 × 10–2 0.1345 2.132 0.09981 9.9625 × 10 –3 2.1240 × 10–2 ∑Z2 0.1129 ∑(Z2t) T √∑ܼଶ CI ∑ሺܼଶݐሻ ∑ ܼଶ ܶ√∑ܼଶ 0.2060 1.8248 0.3360 0.6131 **ADJUSTMENT OF AIRCRAFT NOISE LEVELS FOR THE EFFECTS** OF BACKGROUND NOISE [ETM 3.6] 306(a) Introduction [ETM 3.6.1] The following is provided as guidance material on procedures for adjusting measured aircraft noise levels for the effects of background noise. The presence of background noise during aircraft noise certification tests can influence measured aircraft sound levels and, in some cases, obscure portions of the spectral time history used to obtain EPNL values. Adjustment procedures should include the following components: a) testing to determine which portions of the spectral time history, if any, are obscured; ### Page 98 b) adjustment of unobscured levels to determine the aircraft sound levels that would have been measured in the absence of background noise; and c) replacement or reconstruction of obscured levels by frequency extrapolation, time extrapolation or other means. Definitions of the terms used in this section are provid ed in 306(b). Although some of the terms have generally accepted meanings, the specific meanings as defined apply herein. A detailed step-by-step procedure is presented in 306(c) including equations and descriptions of time and frequency extrapolation methods (see 306(c)(2)(J)). Other procedures may be used provided that they have been approved by the FAA. General considerations that apply to any background noise adjustment procedure are listed in 306(d), including limitations and requirements (see 306(d)(1)) and other special considerations (see 306(d)(2) through 306(d)(4)). 306(b) Definitions [ETM 3.6.2] For the purposes of 306, the following definitions apply: Adjusted level. A valid one-third octave band level which has been adjusted for measurement conditions, including: a) the energy contribution of pre-detection noise; and b) frequency-dependent adjustments such as system frequency response, microphone pressure response and free-field response, and windscreen incidence-dependent insertion loss. Ambient noise. The acoustical noise from sources other than the te st aircraft present at the microphone site during aircraft noise measurements. Ambient noise is one component of background noise. Background noise. The combined noise present in a measurement system from sources other than the test aircraft, which can influence or obscure the aircraft noise levels being measured. Typical elements of background noise include, but are not limited to, ambient noise from sour ces around the microphone s ite, thermal electrical noise generated by components in the measurement system, magnetic flux noise (“tape hiss”) from analog tape recorders and digitization noise caused by quantization error in digital converters. Some elements of background noise, such as digitization noise, can obscure the aircraft noise signal, while others, such as ambient noise, can also contribute energy to the measured aircraft noise signal. Energy-subtraction. Subtraction of one sound pressure level from another, on an energy basis, in the form of the following: 10݃݋݈ଵ଴ሾ10ሺ௅ଵ଴ಲሻ െ1 0ሺ௅ଵ଴ಳሻሿ where L A and LB are two sound pressure levels in decibels, with LB being the value subtracted from LA. Frequency extrapolation. A method for reconstruction of high frequency masked data, based on unmasked data in a lower-frequency one-third octave band from the same spectrum. High frequency bands. The twelve bands from 800 Hz through 10 kHz inclusive (also see “low frequency bands”). Last good band (LGB). In the adjustment methodology presented in 3.6.3, for any aircraft one-third band spectrum, Page 98 ### Page 99 the LGB is the highest frequency unmasked band within the range of 630 Hz to 10 kHz inclusive, below which there are no masked high frequency bands. Low frequency bands. The twelve bands from 50 Hz through 630 Hz inclusive (also see “high frequency bands”). Masked band. Within a single spectrum, any one-third octave band containing a masked level. Masked level. Any one-third octave band level which is less than or equal to the masking criterion for that band. When a level is identified as being masked, the actual level of aircraft noise in that band has been obscured by background noise and cannot be determined. Masked levels can be reconstructed using frequency extrapolation, time extrapolation or other methods. Masking criteria. The spectrum of one-third octave band levels below which measured aircraft sound pressure levels are considered to be masked or obscured by background noise. Masking criteria levels are defined as the greater of: a) pre-detection noise + 3 dB; or b) post-detection noise + 1 dB. Post-detection noise. The minimum levels below which measured noise levels are not considered valid. Usually determined by the baseline of an analysis “window” or by the amplitude non-linearity characteristics of components in the measurement and analysis system. Post-detection noise levels are non-additive (i.e., they do not contribute energy to measured aircraft noise levels). Pre-detection noise. Any noise which can contribute energy to the measured levels of sound produced by the aircraft, including ambient noise present at the microphone site and active instrumentation noise present in the measurement, record/playback and analysis systems. Reconstructed level. A level, calculated by frequency extrapolation, time extrapolation, or by other means, which replaces the measured value for a masked band. Sound attenuation coefficient. The reduction in level of sound within a one-third octave band, in dB per 100 meters, due to the effects of atmospheric absorption of sound. Time extrapolation. A method for reconstruction of high frequency masked data, based on unmasked data in the same one-third octave band, from a different spectrum in the time history. Valid or unmasked band. Within a single spectrum, any one-third octave band containing a valid level. Valid or unmasked level. Any one-third octave band level which exceeds the masking criterion for that band. 306(c) Background Noise Adjustment Procedure [ETM 3.6.3] 306(c)(1) Assumptions [ETM 3.6.3.1] a) A typical aircraft spectru m measured on the ground contains one-third octave band levels which decrease in amplitude with increasing frequency. This characteristic high frequency roll-off is due primarily to the effects of atmospheric absorption. b) A typical electronic instrumentation floor spectrum contains one-third octave band levels which increase in amplitude with increasing frequency. ### Page 100 c) Due to the assumptions cited in a) and b), as the observed frequency is increased within a one-third octave band aircraft spectrum and once a band becomes masked, all subsequent higher frequency bands will also be masked. This allows the implementation of a Last Good Band (LGB) label to identify the frequency band above which the bands in a spectrum are masked. d) If, on occasion, a valid level occurs in a band with higher center frequency than the LGB, its presence will most likely be due to small variations in the pre-detection levels and/or due to levels of the measured aircraft one-third octave band spectrum being close to the levels of the background noise in general, so its energy contribution will not be significant. Note that this assumption is valid only in the absence of significant aircraft-generated tones in the region of masking. Th erefore, the possibility of a level being valid in a band with higher center frequency than the LGB may be ignored. Applicants who prefer to implement algorithms for identifying and handling such situations may do so, but no procedure may be used without prior approval by the FAA. 306(c)(2) Step-by step description [ETM 3.6.3.2] 306(c)(2)(A) Determination of pre-detection noise [ETM 3.6.3.2.1] A time-averaged one-third octave band spectrum of pre-detection noise levels for each test run, or group of runs occurring during a short time period, should be obtained by recording and analyzing ambient noise over a representative period of time (30 seconds or more). Care should be taken to ensure that this “ambient” noise sample reasonably represents that which is present during measured aircraft runs. In recording ambient noise, all gain stages and attenuators should be set as they would be during the aircraft runs in order to ensure that the instrumentation noise is also representative. If multiple gain settings are required for aircraft noise measurements, a separate ambient sample should be recorded at each of the settings used. 306(c)(2)(B) Determination of post-detection noise [ETM 3.6.3.2.2] A one-third octave band spectrum of post-detection noise levels should be determined as a result of testing, or from manufacturer’s specifications, for each measurement/analysis configuration used, including different gain and/or sensitivity settings. These minimum valid levels may be determined on the basis of display limitations (e.g., blanking of the displayed indication when levels fall below a cert ain value), amplitude non-linearity or other non-additive limitations. In cases where more than one component or stage of the measurement/analysis system imposes a set of minimum valid levels, the most restrictive in each one-third octave band should be used. 306(c)(2)(C) Testing of pre-detectio n noise versus post-detection noise [ETM 3.5.3.2.3] The validity of pre-detection noise levels must be established before these levels can be used to adjust valid aircraft noise levels. Any pre-detection noise level that is equal to or less than the post-detection noise level in a particular one-third octave band should be identified as invalid and therefore should not be used in the adjustment procedure. 306(c)(2)(D) Determinati on of masking criteria [ETM 3.6.3.2.4] Once the pre-detection noise and post-detection noise spectra are established, the masking criteria can be identified. For each one-third octave band, compare the valid pre-detection noise level + 3 dB with the post-detection noise level + 1 dB. The highest of these levels is used as the masking criterion for that band. If there is no valid pre-detection noise level for a particular one-third octave band, then th e post-detection noise level + 1 dB is used as the masking criterion for that band. The 3 dB window above pre-detection levels allows for the doubling of energy which could Page 100 ### Page 101 occur if an aircraft noise level were equal to the pre-detection level. The 1 dB window above the post-detection levels allows for a reasonable amount of error in the determination of those levels. 306(c)(2)(E) Identification of masked levels [ETM 3.6.3.2.5] Each spectrum in the aircraft noise time history can be evaluated for masking by comparing the one-third octave band levels against the masking criteria levels. Whenever the aircraft level in a particular band is less than or equal to the associated masking criterion, that aircraft level is considered masked. A record must be kept of which bands in each spectrum are masked. 306(c)(2)(F) Determinati on of Last Good Band [ETM 3.6.3.2.6] For each half-second spectral record, determine the highest frequency unmasked one-third octave band (“Last Good Band ” or “LGB”) by starting at the 630 Hz band and incrementing the band number (i.e., increasing frequency) until a masked band is found. At that point, set LGB for that spectral record equal to the band below the masked band. The lowest frequency band that can be identified as LGB is the 630 Hz band. In other words, if both the 630 Hz band and the 800 Hz band are masked, no reconstruction of masked levels may be performed for that spectrum, and the thirteen bands between 630 Hz and 10 kHz inclusive should be left as is and identified as masked. According to the masking limits specified in 306(d)(2), such a spectrum is not valid for calculation of EPNL. 306(c)(2)(G) Adjustment of valid levels for background noise [ETM 3.6.3.2.7] In each half-second spectrum, for each valid band up to and including LGB, perform an energy-subtraction of the valid pre-detection level from the valid measured level in the aircraft noise time history using: 10݃݋݈ଵ଴ሾ10ሺ ௅ಲ೔ೝ೎ೝೌ೑೟ሻ െ1 0ሺ௅ುೝ೐೏೐೟೐೎೟೔೚೙ሻଵ଴ ଵ଴ ሿ Energy-subtraction should be performed on all valid one-third octave band noise levels. For any one-third octave band where there is no valid pre-detection noise level, no energy-subtraction may be performed (i.e., this adjustment cannot be applied when either the measured aircraft noise time-history level or the pre-detection noise level is masked). 306(c)(2)(H) Adjustment of valid levels for measurement conditions [ETM 3.6.3.2.8] Before any reconstruction can be done for masked levels, valid levels which have been adjusted for the presence of pre-detection noise must also then be adjusted for frequency-dependent adjustments such as system frequency response, microphone pressure response and free-field response, and windscreen incidence-dependent insertion loss. These adjustments cannot be applied to masked levels. 306(c)(2)(I) Reconstruction of low frequency masked bands [ETM 36.3.2.9] In cases where a single masked low frequency one-third octave band occurs between two adjacent valid bands, the masked level can be retained, or the arithmetic average of the adjusted levels of the adjacent valid bands may be used in place of the masked level. If the average is used, the le vel should be categorized as reconstructed. However, if masked low frequency bands are found adjacent to other masked low frequency bands, these masked levels should be retained and remain categorized as ma sked. The procedure presented in this 306(c) does not provide for any other form of reconstruction for masked low frequency bands. ### Page 102 306(c)(2)(J) Reconstruction of leve ls for masked high frequency bands [ETM 3.6.3.2.10] Frequency extrapolation and time extrapolation are the methods used to reconstruct masked one-third octave band levels for bands at frequencies higher than LGB for each spectral record. One-third octa ve band sound attenuation coefficients (either in dB per 100 m or in dB per 1000 ft) must be determined before such reconstruction of masked band levels can be performed. Note that sound emission co ordinates must also be calcu lated for each record before reconstruction is performed since the procedure is dependent on propagation distance. 306(c)(2)(J)(i) Frequency extrapolation method [ETM 3.6.3.2.10.1] For a spectrum where the LGB is located at or above the 2 kHz one-third octave band, the frequency extrapolation method is used. This method reconstructs masked high frequency bands starting with the level associated with LGB in the same spectrum. The levels for all bands at higher frequencies than LGB must be reconstructed using this method. Any frequency-extrapolated levels should be categorized as reconstructed. Reconstruct the level for the masked bands using the following equation: ܴܵ௞ 60 ܴܵ௞ 60 ܴܵ௞ ܴܵ௞ܮ௜,௞ ൌܮ ௝,௞ ൅ߙ ௝ 100 െߙ ௝ோாி 100 ൅ 20݃݋݈ଵ଴ 60 ൅ߙ ௝ோாி 100 െߙ ௝ 100 ൅ 20݃݋݈ଵ଴ 60 which can be reduced to: ௌோೖ ଺଴ݔܮ௜,௞ ൌܮ ௝,௞ ൅ሾ ߙ௝ െߙ ௜ሿ ଵ଴଴ ൅ሾ ߙ௜ோாி െߙ ௝ோாிሿ ଵ଴଴, where: — i is the masked band to be extrapolated; — k is the record of interest; — j is the LGB in record k; — Lxi,k is the frequency-extrapolated level in dB for masked band i and spectral record k; — Lj,k is the level for LGB in record k after all test-day adjustments have been applied, including predetection noise energy-subtraction, system and microphone adjustments, etc.; — αj is the test-day sound attenuation coefficient (dB per 100 m) for LGB; αi is the test-day sound attenuation coefficient (dB per 100 m) for band i; αjREF is the reference 77°F (25°C), 70 percent relative humidity (RH) sound attenuation coefficient (dB per 100 m) for LGB; αiREF is the reference 77°F (25°C), 70 percent RH) sound attenuation coefficient (dB per 100 m) for masked band i; and SRk is the slant range or acoustic propagation distance in meters at the time of sound emission for spectral record k, between the aircraft and the microphone. This procedure is based on the assumption that the aircraft spectrum is “flat” (i.e., all high frequency band levels are equal) at a distance of 197 ft (60 m) under reference conditions 77°F (25°C), 70 percent RH). The process can be conceptualized by means of the following steps: ### Page 103 a) the level for band j, the highest frequency unmasked band in spectral record k, which has already been adjusted for measurement conditions, is adjusted for test-day propagation effects to obtain the source level and then adjusted using reference propagation effects to the197 ft (60 m) distance from the source; b) this level is then assigned as the level for all high frequency masked bands (i.e., band i, band i + 1, etc.) at a distance of 197 ft (60 m); c) a new source level is determined for each masked high frequency band by removing the associated reference-day propagation effects; and d) the extrapolated level that would have been measured on the ground, in the absence of background noise, is determined for each masked high frequency band by adding the test-day propagation effects to each of the source levels determined in c) above. 306(c)(2)(J)(ii) Time extrapolation method [ETM 3.6.3.2.10.2] For a spectrum where LGB occurs at or between the 630 Hz one-third octave band and the 1.6 kHz band, use the time extrapolation method. This method reconstructs a masked band in a spectrum from the closest spectral record (i.e., closest in time) for which that band is valid. The levels fo r all one-third octave bands with frequencies greater than that of LGB must be reconstructed using this time extr apolation method. Any time-extrapolated levels should be categorized as reconstructed. Reconstruct the levels for the masked bands by using the following equation: ሾܴܵ௠ܮ௜,௞ ൌܮ ௜,௠ ൅ߙ ௝ 100ሿ ൅ 20݃݋݈ଵ଴ሾܴܵ௠ሿ100 െ ܴܵ௞ ܴܵ௞ where: — L xi,k is the time-extrapolated level in dB for masked band i and spectral record k; — Li,m is the adjusted level in dB for band i in spectral record m, which is the nearest record in time to record k in which band i contains a valid level; — SRm is the slant range or acoustic propagation distance in meters at the time of sound emission for spectral record m, between the aircraft and the microphone; — SRk is the slant range or acoustic propagation distance in meters at the time of sound emission for spectral record k, between the aircraft and the microphone; and — αiis the test-day sound attenuation coefficient (dB per 100 m) for band i. This procedure is based on the assumption that the aircraft spectrum is omnidirectional during the aircraft pass-by. 306(c)(2)(K) Handling of spectra af ter reconstruction of masked bands [ETM 3.6.3.2.11] After reconstruction of masked data has been performed, the background noise adjustment procedure is complete. The adjusted as-measured data set, comprised of adjusted levels, reconstructed levels, and possibly some masked levels, is next used to obtain the test-day PNLT time history desc ribed in A36.4.3 of part 36. The identification of masked data should be kept accessible for use during the tone co rrection procedure, since any tone correction which results from the adjustment for background noise may be eliminated from the process of identifying the maximum tone within a spectrum. When this background noise adjustment procedure is used, the band identified as LGB should be treated Page 103 ### Page 104 as the last band of the tone correction calculation in the manner prescribed for the 10 kHz band in A36.4.3.1 of part 36, including the calculation of a new slope for band LGB + 1 that equals the slope at LGB (i.e., s ′(LGB + 1, k) = s′(LGB,k)) in Step 5 of the tone correction procedure. 306(d) General Considerations [ETM 3.6.4] 306(d)(1) Limitations and requirements for an y background noise adjustment procedure [ETM 3.6.4.1] Any method of adjusting for the effects of background noise must be approved by the FAA before it is used. The adjustment procedure presented in 306(c)(2) includes applicable limitations and requirements. Those limitations and requirements which apply to all methodologies are described as follows. The applicant must be able to demonstrate by means of narrow-band analysis or other methods that no significant aircraft-generated tones occur in the masked one-third octave bands during the EPNL duration. Neither frequency-dependent adjustments nor energy-subtraction of pre-detection levels can be applied to masked data. When consecutive one-third octave bands in the range of 2.5 kHz to 10 kHz inclusiv e are masked, and when no consecutive bands are masked in the region of 800 Hz to 2 kHz inclusive, frequency extrapolation, as described in 306(c)(2)(J)(i), must be performed on all consecutive masked bands with nominal frequencies greater than 2 kHz. When consecutive one-third octave bands in the range of 800 Hz to 2 kHz inclusive are masked, time extrapolation, as described in 306(c)(2)(J)(ii), must be performed on all consecutive masked bands with nominal frequencies greater than 630 Hz. In cases where a single masked one-third octave band occurs between two adjacent valid bands, the levels of the adjacent adjusted bands may be arithmetically averaged and the averaged level used in place of the masked level. If the masked level is retained it must be included when counting the masked levels in the procedure described in 306(d)(2) 306(d)(2) Rejection of spectra due to masking [ETM 3.6.4.2] A spectrum becomes invalid if the following conditions prevail: a) if, after any reconstruction of masked bands, more than four one-third octave bands retain masked values; b) for records within one second of the record associated with the PNLTmax spectrum (i.e., five half-second data records) if: 1) more than four high frequency bands require reconstruction; or 2) the LGB is located at or below the 3150 Hz one-third octave band when the example background noise adjustment procedure presented in 306(c)(2) is used. Note.— If an invalid spectrum occurs within the 10 dB-down period, the aircraft test run is invalid and cannot be used for aircraft noise certification purposes. 305(d)(3) Special tone correction considerations due to masking [ETM 3.6.4.3] Page 104 ### Page 105 307 When the maximum tone correction for a one-third octave band spectrum occurs at a masked or reconstructed band, the tone correction for that spectrum cannot simply be set to zero. The maximum tone correction for the spectrum must be computed, taking masked or reconstructed levels into consideration. Any tone correction resulting from the adjustment for background noise may be eliminated by either one of the following two methods, as appropriate: a) when the example background noise adjustment proce dure presented in 306(c)(2) is used or, specifically, when all of the high frequency bands in a spectrum are masked for frequencies beyond a certain band (i.e., “LGB”) the band labeled as LGB should be treated as the last band of the tone correction calculation, in the manner prescribed for the 10 kHz band (band number 24) in A36.4.3.1 of part 36, including calculation of a new slope for the band above LGB that equals the slope of the band at LGB (i.e., s′(LGB + 1,k) = s′(LGB,k)) in Step 5 of the tone correction procedure; or b) for tone corrections that occur at one-third octave bands that are masked or reconstructed, set F equal to zero in Step 9 of the tone correction procedure, and recalculate the maximum tone correction for that spectrum. Note.— All band levels within a spectru m, whether adjusted, reconstructed or masked, must be included in the computation of the PNL value for that spectrum. 306(d)(4) Handling of masked data in reference conditions data sets [ETM 3.6.4.4] For any one-third octave band spectrum adjusted to reference conditions, all bands, including those containing masked levels or reconstructed levels, includin g values less than 0 dB, must be adju sted for differences between test and reference conditions (i.e., atmospheric absorption and spherical spreading). The special tone correction considerations listed in 306(c)(2) apply both to test and reference data sets. Noise Reduction Systems [ETM 3.7] An aircraft can employ noise reduction systems that chan ge its configuration or operating condition to reduce noise, or implement devices or subsystems that directly reduce or counteract noise emissions. Two categories, variable noise reduction systems (VNRS) and selectable noise reduction systems (SNRS), have been defined to address differences in activation/actuation for these systems. General guidan ce on noise certification of ai rcraft equipped with these systems is provided below. 307(a) Variable Noise Reduction Systems [ETM 3.7.1] A VNRS is an integral design feature, or subsystem, of an aircraft that automatically changes the configuration or operating condition of the aircraft to reduce noise. Note 1.— If pilot action is necessary to activate, i.e., select the use of, an automatically controlled noise reduction system or if a pilot can deactivate (deselect) an automatically controlled noise reduction system, such a system is not considered a VNRS. Note 2.— Aircraft can incorporate variable systems, primarily intended to improve performance, reduce engine emissions and/or increase safety, that may also affect noise. Such aircraft can be noise certificated using the guidance provided for aircraft with VNRS. For such changes to existing type designs, the guidelines provided in 101(d) for “no acoustical change” are applicable. For a VNRS-equipped aircraft, the VNRS characteristics may prevent flight from being conducted in accordance with the associated reference procedure(s) in part 36. In such cases, the reference procedures for noise certification of an aircraft with a VNRS should depart from those specified in part 36 only to the extent required by those design Page 105 ### Page 106 characteristics that cause the departure and should be approved by the FAA after the rulemaking process for an deviation from part 36 is completed (see 201(a)(10) paragraph 2 of this AC). The impact of a VNRS on nois e certification of an aircraft can extend beyo nd deviations from the part 36 reference procedures. A plan for noise certificatio n of a VNRS-equipped aircraft should take into consideration three key elements, namely: a) the necessity, if any, to depart from the part 36 procedures; b) the adaptation/modification of test proced ures to ensure compliance with part 36; and c) the applicability of existing procedures in part 36 for adjusting the measured data to reference conditions. Experience to date has shown that one or more of these elements may be interrelated, requiring detailed consideration of all three elements in devising an acceptable plan for noise certification. 307(a)(1) Reference procedures [ETM 3.7.1.1] The part 36 reference procedures typically utilize consta nt flight path and operational parameters. A VNRS can, however, result in non-constant reference flight paths and/or non-constant operational parameters such as non-constant rates of climb and/or non-constant engine/propeller/rotor speeds, respectively, that compel departure from the reference procedures. In addition to reducing noise emissions, a VNRS may, and typically does, impact aircraft performance during a noise certification reference procedure. In some cases, this impact can be indirect via another affected performance parameter. Both direct and indirect impacts on aircraft performance should be addressed in defining any departure necessary from the reference procedures in part 36 to accommodate a VNRS. Actuation of a VNRS can be a function of one or more ope rational conditions such as airspeed, ground speed, height above ground level, density altitude, pressure altitude and ambient temperature. Beginning and end points on the reference flight path for any transition triggered by a VNRS should be determined using the reference test and meteorological conditions. 307(a)(2) Test conditions and procedures [ETM 3.7.1.2] When a VNRS results in a non-constant reference flight path for the aircraft, the flight path tolerances (height and lateral deviation limits) specified in part 36 for the corresponding constant reference procedure should be applied, subject to approval by the FAA. Similarly, when a VNRS results in a non-constant operational parameter for the aircraft, a reference schedule for the affected operating parameter should be defined along the reference flight path, and the test tolerances permitted by part 36 for that parameter should be applied to the reference schedule, subject to approval by the FAA. 307(a)(3) Adjustments to measured noise data [ETM 3.7.1.3] Adjustments to measured data in part 36 are based on constant reference procedures. A VNRS can, however, result in a non-constant reference procedure(s) that in turn impacts the adjustments to measured data that account for test deviations from reference flight profiles and test conditions. The adjustments to measured data specified in part 36 should be modified only as necessary to account for any departures from the reference procedures in part 36. In many cases, only minor changes to data-processing software that do not affect the adjustment procedures will be needed. Any modifications, including software revisions, of the adjustments to measured data specified in part 36 are subject to approval by the FAA. ### Page 107 307(a)(4) Guidance for specific VNRS [ETM 3.7.1.4] Specific guidance for VNRS technologies will typically be developed as these technologies are developed and implemented in aircraft designs. Cross references to the appropriate sections of this Advisory Circular for the VNRS technologies for which specific guidance has been generally accepted by FAA are provided in Table 8. Table 8. Cross references to spec ific guidelines in this ACfor VNRS Variable noise reduction Specific guidelines provided in system (VNRS) this AC Variable rotor speed Chapter 4, 401(h), H36.3: helicopters H36.103(b) [Takeoff flight test procedures], paragraph 8) 307(b) Selectable Noise Reduction Systems [ETM 3.7.2] (Reserved) Note.— The guidance provided in 307 addresses VNRS only and, by inference, defines as selectable all noise reduction systems that do not satisfy the requirements for classification as VNRS. Guidance specific to SNRS, including a definition specific to SNRS, is not yet provided. 308 CALCULATION OF THE SPEED OF SOUND [ETM 3.8] For the purposes of noise certification the value of the speed of sound, c, is calculated from the equation taken from ISO 9613-1:1993(E): c = 1125.9 (T/T0)½ ft/s, or c = 343.2 (T/T0)½ m/s where T0 = 293.15 K and T is the absolute ambient air temperature in kelvins. Note.— At the noise certification reference temperature of 25°C, T = 298.15 K, and c therefore equals 1135.5 ft/s (346.1 m/s). **309 REFERENCE TABLES USED IN THE MANUAL CALCULATION OF** OF EFFECTIVE PERCEIVED NOISE LEVEL [ETM 3.9] Table 9 and Table 10 and Figure 7 contain information useful for the manual calculation of EPNL. Such manual calculations are often used to verify the accuracy of computer programs used for calculating certification noise levels. ### Page 108 5 10 15 20 25 30 35 40 45 50 55 10/12/2017 AC 36-4D Table 9. Perceived noisiness (noys) as a function of sound pressure level (One-third octave band center frequencies (Hz)) SPL 50 63 80 100 125 160 200 250 315 400 500 630 800 1000 1250 1600 2000 2500 3150 4000 5000 6300 8000 10000 4 0.10 0.10 0.11 0.10 6 0.11 0.12 0.11 0.10 7 0.12 0.14 0.13 0.11 8 0.14 0.16 0.14 0.13 9 0.10 0.16 0.17 0.16 0.14 0.11 0.17 0.19 0.18 0.16 0.10 11 0.13 0.19 0.22 0.21 0.18 0.12 12 0.10 0.14 0.22 0.24 0.24 0.21 0.14 13 0.11 0.16 0.24 0.27 0.27 0.24 0.16 14 0.13 0.18 0.27 0.30 0.30 0.27 0.19 0.10 0.14 0.21 0.30 0.33 0.33 0.30 0.22 16 0.10 0.10 0.10 0.10 0.10 0.11 0.16 0.24 0.33 0.35 0.35 0.33 0.26 17 0.11 0.11 0.11 0.11 0.11 0.13 0.18 0.27 0.35 0.38 0.38 0.35 0.30 0.10 18 0.10 0.13 0.13 0.13 0.13 0.13 0.15 0.21 0.30 0.38 0.41 0.41 0.38 0.33 0.12 19 0.11 0.14 0.14 0.14 0.14 0.14 0.17 0.24 0.33 0.41 0.45 0.45 0.41 0.36 0.14 0.13 0.16 0.16 0.16 0.16 0.16 0.20 0.27 0.36 0.45 0.49 0.49 0.45 0.39 0.17 21 0.10 0.14 0.18 0.18 0.18 0.18 0.18 0.23 0.30 0.39 0.49 0.53 0.53 0.49 0.42 0.21 0.10 22 0.11 0.16 0.21 0.21 0.21 0.21 0.21 0.26 0.33 0.42 0.53 0.57 0.57 0.53 0.46 0.25 0.11 23 0.13 0.18 0.24 0.24 0.24 0.24 0.24 0.30 0.36 0.46 0.57 0.62 0.62 0.57 0.50 0.30 0.13 24 0.10 0.14 0.21 0.27 0.27 0.27 0.27 0.27 0.33 0.40 0.50 0.62 0.67 0.67 0.62 0.55 0.33 0.15 0.11 0.16 0.24 0.30 0.30 0.30 0.30 0.30 0.35 0.43 0.55 0.67 0.73 0.73 0.67 0.60 0.36 0.17 26 0.13 0.18 0.27 0.33 0.33 0.33 0.33 0.33 0.38 0.48 0.60 0.73 0.79 0.79 0.73 0.65 0.39 0.20 27 0.10 0.14 0.21 0.30 0.35 0.35 0.35 0.35 0.35 0.41 0.52 0.65 0.79 0.85 0.85 0.79 0.71 0.42 0.23 28 0.11 0.16 0.24 0.33 0.38 0.38 0.38 0.38 0.38 0.45 0.57 0.71 0.85 0.92 0.92 0.85 0.77 0.46 0.26 29 0.13 0.18 0.27 0.35 0.41 0.41 0.41 0.41 0.41 0.49 0.63 0.77 0.92 1.00 1.00 0.92 0.84 0.50 0.30 0.10 0.14 0.21 0.30 0.38 0.45 0.45 0.45 0.45 0.45 0.53 0.69 0.84 1.00 1.07 1.07 1.00 0.92 0.55 0.33 31 0.11 0.16 0.24 0.33 0.41 0.49 0.49 0.49 0.49 0.49 0.57 0.76 0.93 1.07 1.15 1.15 1.07 1.00 0.60 0.37 32 0.13 0.18 0.27 0.36 0.45 0.53 0.53 0.53 0.53 0.53 0.62 0.83 1.00 1.15 1.23 1.23 1.15 1.07 0.65 0.41 33 0.14 0.21 0.30 0.39 0.49 0.57 0.57 0.57 0.57 0.57 0.67 0.91 1.07 1.23 1.32 1.32 1.23 1.15 0.71 0.45 34 0.10 0.16 0.24 0.33 0.42 0.53 0.62 0.62 0.62 0.62 0.62 0.73 1.00 1.15 1.32 1.41 1.41 1.32 1.23 0.77 0.50 0.11 0.18 0.27 0.36 0.46 0.57 0.67 0.67 0.67 0.67 0.67 0.79 1.07 1.23 1.41 1.51 1.51 1.41 1.32 0.84 0.55 36 0.13 0.21 0.30 0.40 0.50 0.62 0.73 0.73 0.73 0.73 0.73 0.85 1.15 1.32 1.51 1.62 1.62 1.51 1.41 0.92 0.61 37 0.15 0.24 0.33 0.43 0.55 0.67 0.79 0.79 0.79 0.79 0.79 0.92 1.23 1.41 1.62 1.74 1.74 1.62 1.51 1.00 0.67 38 0.17 0.27 0.37 0.48 0.60 0.73 0.85 0.85 0.85 0.85 0.85 1.00 1.32 1.51 1.74 1.86 1.86 1.74 1.62 1.10 0.74 39 0.10 0.20 0.30 0.41 0.52 0.65 0.79 0.92 0.92 0.92 0.92 0.92 1.07 1.41 1.62 1.86 1.99 1.99 1.86 1.74 1.21 0.82 0.12 0.23 0.33 0.45 0.57 0.71 0.85 1.00 1.00 1.00 1.00 1.00 1.15 1.51 1.74 1.99 2.14 2.14 1.99 1.86 1.34 0.90 41 0.14 0.26 0.37 0.50 0.63 0.77 0.92 1.07 1.07 1.07 1.07 1.07 1.23 1.62 1.86 2.14 2.29 2.29 2.14 1.99 1.48 1.00 42 0.16 0.30 0.41 0.55 0.69 0.84 1.00 1.15 1.15 1.15 1.15 1.15 1.32 1.74 1.99 2.29 2.45 2.45 2.29 2.14 1.63 1.10 43 0.19 0.33 0.45 0.61 0.76 0.92 1.07 1.23 1.23 1.23 1.23 1.23 1.41 1.86 2.14 2.45 2.63 2.63 2.45 2.29 1.79 1.21 44 0.10 0.22 0.37 0.50 0.67 0.83 1.00 1.15 1.32 1.32 1.32 1.32 1.32 1.52 1.99 2.29 2.63 2.81 2.81 2.63 2.45 1.99 1.34 0.12 0.26 0.42 0.55 0.74 0.91 1.08 1.24 1.41 1.41 1.41 1.41 1.41 1.62 2.14 2.45 2.81 3.02 3.02 2.81 2.63 2.14 1.48 46 0.14 0.30 0.46 0.61 0.82 1.00 1.16 1.33 1.52 1.52 1.52 1.52 1.52 1.74 2.29 2.63 3.02 3.23 3.23 3.02 2.81 2.29 1.63 47 0.16 0.34 0.52 0.67 0.90 1.08 1.25 1.42 1.62 1.62 1.62 1.62 1.62 1.87 2.45 2.81 3.23 3.46 3.46 3.23 3.02 2.45 1.79 48 0.19 0.38 0.58 0.74 1.00 1.17 1.34 1.53 1.74 1.74 1.74 1.74 1.74 2.00 2.63 3.02 3.46 3.71 3.71 3.46 3.23 2.63 1.98 49 0.10 0.22 0.43 0.65 0.82 1.08 1.26 1.45 1.64 1.87 1.87 1.87 1.87 1.87 2.14 2.81 3.23 3.71 3.97 3.97 3.71 3.46 2.81 2.18 0.12 0.26 0.49 0.72 0.90 1.17 1.36 1.56 1.76 2.00 2.00 2.00 2.00 2.00 2.30 3.02 3.46 3.97 4.26 4.26 3.97 3.71 3.02 2.40 51 0.14 0.30 0.55 0.80 1.00 1.26 1.47 1.68 1.89 2.14 2.14 2.14 2.14 2.14 2.46 3.23 3.71 4.26 4.56 4.56 4.26 3.97 3.23 2.63 52 0.17 0.34 0.62 0.90 1.08 1.36 1.58 1.80 2.03 2.30 2.30 2.30 2.30 2.30 2.64 3.46 3.97 4.56 4.89 4.89 4.56 4.26 3.46 2.81 53 0.21 0.39 0.70 1.00 1.18 1.47 1.71 1.94 2.17 2.46 2.46 2.46 2.46 2.46 2.83 3.71 4.26 4.89 5.24 5.24 4.89 4.56 3.71 3.02 54 0.25 0.45 0.79 1.09 1.28 1.58 1.85 2.09 2.33 2.64 2.64 2.64 2.64 2.64 3.03 3.97 4.56 5.24 5.61 5.61 5.24 4.89 3.97 3.23 0.30 0.51 0.89 1.15 1.35 1.71 2.00 2.25 2.50 2.83 2.83 2.83 2.83 2.83 3.25 4.26 4.89 5.61 6.01 6.01 5.61 5.24 4.26 3.46 56 0.34 0.59 1.00 1.29 1.50 1.85 2.15 2.42 2.69 3.03 3.03 3.03 3.03 3.03 3.48 4.56 5.24 6.01 6.44 6.44 6.01 5.61 4.56 3.71 57 0.39 0.67 1.09 1.40 1.63 2.00 2.33 2.61 2.88 3.25 3.25 3.25 3.25 3.25 3.73 4.89 5.61 6.44 6.90 6.90 6.44 6.01 4.89 3.97 58 0.45 0.77 1.18 1.53 1.77 2.15 2.51 2.81 3.10 3.48 3.48 3.48 3.48 3.48 4.00 5.24 6.01 6.90 7.39 7.39 6.90 6.44 5.24 4.26 59 0.51 0.87 1.29 1.66 1.92 2.33 2.71 3.03 3.32 3.73 3.73 3.73 3.73 3.73 4.29 5.61 6.44 7.39 7.92 7.92 7.39 6.90 5.61 4.56 Page 108 ### Page 109 SPL 50 63 80 100 125 160 200 250 315 400 500 630 800 1000 1250 1600 2000 2500 3150 4000 5000 6300 8000 10000 60 0.59 1.00 1.40 1.81 2.08 2.51 2.93 3.26 3.57 4.00 4.00 4.00 4.00 4.00 4.59 6.01 6.90 7.92 8.49 8.49 7.92 7.39 6.01 4.89 61 0.67 1.10 1.53 1.97 2.26 2.71 3.16 3.51 3.83 4.29 4.29 4.29 4.29 4.29 4.92 6.44 7.39 8.49 9.09 9.09 8.49 7.92 6.44 5.24 62 0.77 1.21 1.66 2.15 2.45 2.93 3.41 3.78 4.11 4.59 4.59 4.59 4.59 4.59 5.28 6.90 7.92 9.09 9.74 9.74 9.09 8.49 6.90 5.61 63 0.87 1.32 1.81 2.34 2.65 3.16 3.69 4.06 4.41 4.92 4.92 4.92 4.92 4.92 5.66 7.39 8.49 9.74 10.4 10.4 9.74 9.09 7.39 6.01 64 1.00 1.45 1.97 2.54 2.88 3.41 3.98 4.38 4.73 5.28 5.28 5.28 5.28 5.28 6.06 7.92 9.09 10.4 11.2 11.2 10.4 9.74 7.92 6.44 65 1.11 1.60 2.15 2.77 3.12 3.69 4.30 4.71 5.08 5.66 5.66 5.66 5.66 5.66 6.50 8.49 9.74 11.2 12.0 12.0 11.2 10.4 8.49 6.90 66 1.22 1.75 2.34 3.01 3.39 3.98 4.64 5.07 5.45 6.06 6.06 6.06 6.06 6.06 6.96 9.09 10.4 12.0 12.8 12.8 12.0 11.2 9.09 7.39 67 1.35 1.92 2.54 3.28 3.68 4.30 5.01 5.46 5.85 6.50 6.50 6.50 6.50 6.50 7.46 9.74 11.2 12.8 13.8 13.8 12.8 12.0 9.74 7.92 68 1.49 2.11 2.77 3.57 3.99 4.64 5.41 5.88 6.27 6.96 6.96 6.96 6.96 6.96 8.00 10.4 12.0 13.8 14.7 14.7 13.8 12.8 10.4 8.49 69 1.65 2.32 3.01 3.88 4.33 5.01 5.84 6.33 6.73 7.46 7.46 7.46 7.46 7.46 8.57 11.2 12.8 14.7 15.8 15.8 14.7 13.8 11.2 9.09 70 1.82 2.55 3.28 4.23 4.69 5.41 6.31 6.81 7.23 8.00 8.00 8.00 8.00 8.00 9.19 12.0 13.8 15.8 16.9 16.9 15.8 14.7 12.0 9.74 71 2.02 2.79 3.57 4.60 5.09 5.84 6.81 7.33 7.75 8.57 8.57 8.57 8.57 8.57 9.85 12.8 14.7 16.9 18.1 18.1 16.9 15.8 12.8 10.4 72 2.23 3.07 3.88 5.01 5.52 6.31 7.36 7.90 8.32 9.19 9.19 9.19 9.19 9.19 10.6 13.8 15.8 18.1 19.4 19.4 18.1 16.9 13.8 11.2 73 2.46 3.37 4.23 5.45 5.99 6.81 7.94 8.50 8.93 9.85 9.85 9.85 9.85 9.85 11.3 14.7 16.9 19.4 20.8 20.8 19.4 18.1 14.7 12.0 74 2.72 3.70 4.60 5.94 6.50 7.36 8.57 9.15 9.59 10.6 10.6 10.6 10.6 10.6 12.1 15.8 18.1 20.8 22.3 22.3 20.8 19.4 15.8 12.8 75 3.01 4.06 5.01 6.46 7.05 7.94 9.19 9.85 10.3 11.3 11.3 11.3 11.3 11.3 13.0 16.9 19.4 22.3 23.9 23.9 22.3 20.8 16.9 13.8 76 3.32 4.46 5.45 7.03 7.65 8.57 9.85 10.6 11.0 12.1 12.1 12.1 12.1 12.1 13.9 18.1 20.8 23.9 25.6 25.6 23.9 22.3 18.1 14.7 77 3.67 4.89 5.94 7.66 8.29 9.19 10.6 11.3 11.8 13.0 13.0 13.0 13.0 13.0 14.9 19.4 22.3 25.6 27.4 27.4 25.6 23.9 19.4 15.8 78 4.06 5.37 6.46 8.33 9.00 9.85 11.3 12.1 12.7 13.9 13.9 13.9 13.9 13.9 16.0 20.8 23.9 27.4 29.4 29.4 27.4 25.6 20.8 16.9 79 4.49 5.90 7.03 9.07 9.76 10.6 12.1 13.0 13.6 14.9 14.9 14.9 14.9 14.9 17.1 22.3 25.6 29.4 31.5 31.5 29.4 27.4 22.3 18.1 80 4.96 6.48 7.66 9.85 10.6 11.3 13.0 13.9 14.6 16.0 16.0 16.0 16.0 16.0 18.4 23.9 27.4 31.5 33.7 33.7 31.5 29.4 23.9 19.4 81 5.48 7.11 8.33 10.6 11.3 12.1 13.9 14.9 15.7 17.1 17.1 17.1 17.1 17.1 19.7 25.6 29.4 33.7 36.1 36.1 33.7 31.5 25.6 20.8 82 6.06 7.81 9.07 11.3 12.1 13.0 14.9 16.0 16.9 18.4 18.4 18.4 18.4 18.4 21.1 27.4 31.5 36.1 38.7 38.7 36.1 33.7 27.4 22.3 83 6.70 8.57 9.87 12.1 13.0 13.9 16.0 17.1 18.1 19.7 19.7 19.7 19.7 19.7 22.6 29.4 33.7 38.7 41.5 41.5 38.7 36.1 29.4 23.9 84 7.41 9.41 10.7 13.0 13.9 14.9 17.1 18.4 19.4 21.1 21.1 21.1 21.1 21.1 24.3 31.5 36.1 41.5 44.4 44.4 41.5 38.7 31.5 25.6 85 8.19 10.3 11.7 13.9 14.9 16.0 18.4 19.7 20.8 22.6 22.6 22.6 22.6 22.6 26.0 33.7 38.7 44.4 47.6 47.6 44.4 41.5 33.7 27.4 86 9.95 11.3 12.7 14.9 16.0 17.1 19.7 21.1 22.4 24.3 24.3 24.3 24.3 24.3 27.9 36.1 41.5 47.6 51.0 51.0 47.6 44.4 36.1 29.4 87 10.0 12.1 13.9 16.0 17.1 18.4 21.1 22.6 24.0 26.0 26.0 26.0 26.0 26.0 29.0 38.7 44.4 51.0 54.7 54.7 51.0 47.6 38.7 31.5 88 11.1 13.0 14.9 17.1 18.4 19.7 22.6 24.3 25.8 27.9 27.9 27.9 27.9 27.9 32.0 41.5 47.6 54.7 58.6 58.6 54.7 51.0 41.5 33.7 89 12.2 13.9 16.0 18.4 19.7 21.1 24.3 26.0 27.7 29.9 29.9 29.9 29.9 29.9 34.3 44.4 51.0 58.6 62.7 62.7 58.6 54.7 44.4 36.1 90 13.5 14.9 17.1 19.7 21.1 22.6 26.0 27.9 29.7 32.0 32.0 32.0 32.0 32.0 36.8 47.6 54.7 62.7 67.2 67.2 62.7 58.6 47.6 38.7 91 14.9 16.0 18.4 21.1 22.6 24.3 27.9 29.9 31.8 34.3 34.3 34.3 34.3 34.3 39.4 51.0 58.6 67.2 72.0 72.0 67.2 62.7 51.0 41.5 92 16.0 17.1 19.7 22.6 24.3 26.0 29.9 32.0 34.2 36.8 36.8 36.8 36.8 36.8 42.2 54.7 62.7 72.0 77.2 77.2 72.0 67.2 54.7 44.4 93 17.1 18.4 21.1 24.3 26.0 27.9 32.0 34.3 36.7 39.4 39.4 39.4 39.4 39.4 45.3 58.6 67.2 77.2 82.7 82.7 77.2 72.0 58.6 47.6 94 18.4 19.7 22.6 26.0 27.9 29.9 34.3 36.8 39.4 42.2 42.2 42.2 42.2 42.2 48.5 62.7 72.0 82.7 88.6 88.6 82.7 77.2 62.7 51.0 95 19.7 21.1 24.3 27.9 29.9 32.0 36.8 39.4 42.2 45.3 45.3 45.3 45.3 45.3 52.0 67.2 77.2 88.6 94.9 94.9 88.6 82.7 67.2 54.7 96 21.1 22.6 26.0 29.9 32.0 34.3 39.4 42.2 45.3 48.5 48.5 48.5 48.5 48.5 55.7 72.0 82.7 94.9 102 102 94.9 88.6 72.0 58.6 97 22.6 24.3 27.9 32.0 34.3 36.8 42.2 45.3 48.5 52.0 52.0 52.0 52.0 52.0 59.7 77.2 88.6 102 109 109 102 94.9 77.2 62.7 98 24.3 26.0 29.9 34.3 36.8 39.4 45.3 48.5 52.0 55.7 55.7 55.7 55.7 55.7 64.0 82.7 94.9 109 117 117 109 102 82.7 67.2 99 26.0 27.9 32.0 36.8 39.4 42.2 48.5 52.0 55.7 59.7 59.7 59.7 59.7 59.7 68.6 88.6 102 117 125 125 117 109 88.6 72.0 100 27.9 29.9 34.3 39.4 42.2 45.3 52.0 55.7 59.7 64.0 64.0 64.0 64.0 64.0 73.5 94.9 109 125 134 134 125 117 94.9 77.2 101 29.9 32.0 36.8 42.2 45.3 48.5 55.7 59.7 64.0 68.6 68.6 68.6 68.6 68.6 78.8 102 117 134 144 144 134 125 102 82.7 102 32.0 34.3 39.4 45.3 48.5 52.0 59.7 64.0 68.6 73.5 73.5 73.5 73.5 73.5 84.4 109 125 144 154 154 144 134 109 88.6 103 34.3 36.8 42.2 48.5 52.0 55.7 64.0 68.6 73.5 78.8 78.8 78.8 78.8 78.8 90.5 117 134 154 165 165 154 144 117 94.9 104 36.8 39.4 45.3 52.0 55.7 59.7 68.6 73.5 78.8 84.4 84.4 84.4 84.4 84.4 97.0 125 144 165 177 177 165 154 125 102 105 39.4 42.2 48.5 55.7 59.7 64.0 73.5 78.8 84.4 90.5 90.5 90.5 90.5 90.5 104 134 154 177 189 189 177 165 134 109 106 42.2 45.3 52.0 59.7 64.0 68.6 78.8 84.4 90.5 97.0 97.0 97.0 97.0 97.0 111 144 165 189 203 203 189 177 144 117 107 45.3 48.5 55.7 64.0 68.6 73.5 84.4 90.5 97.0 104 104 104 104 104 119 154 177 203 217 217 203 189 154 125 108 48.5 52.0 59.7 68.6 73.5 78.8 90.5 97.0 104 111 111 111 111 111 128 165 189 217 233 233 217 203 165 134 109 52.0 55.7 64.0 73.5 78.8 84.4 97.0 104 111 119 119 119 119 119 137 177 203 233 249 249 233 217 177 144 110 55.7 59.7 68.6 78.8 84.4 90.5 104 111 119 128 128 128 128 128 147 189 217 249 267 267 249 233 189 154 111 59.7 64.0 73.5 84.4 90.5 97.0 111 119 128 137 137 137 137 137 158 203 233 267 286 286 267 249 203 165 112 64.0 68.6 78.8 90.5 97.0 104 119 128 137 147 147 147 147 147 169 217 249 286 307 307 286 267 217 177 113 68.6 73.5 84.4 97.0 104 111 128 137 147 158 158 158 158 158 181 233 267 307 329 329 307 286 233 189 114 73.5 78.8 90.5 104 111 119 137 147 158 169 169 169 169 169 194 249 286 329 352 352 329 307 249 203 115 78.8 84.4 97.0 111 119 128 147 158 169 181 181 181 181 181 208 267 307 352 377 377 352 329 267 217 116 84.4 90.5 104 119 128 137 158 169 181 194 194 194 194 194 223 286 329 377 404 404 377 352 286 233 117 90.5 97.0 111 128 137 147 169 181 194 208 208 208 208 208 239 307 352 404 433 433 404 377 307 249 118 97.0 104 119 137 147 158 181 194 208 223 223 223 223 223 256 329 377 433 464 464 433 404 329 267 119 104 111 128 147 158 169 194 208 223 239 239 239 239 239 274 352 404 464 497 497 464 433 352 286 120 111 119 137 158 169 181 208 223 239 256 256 256 256 256 294 377 433 497 533 533 497 464 377 307 121 119 128 147 169 181 194 223 239 256 274 274 274 274 274 315 404 464 533 571 571 533 497 404 329 122 128 137 158 181 194 208 239 256 274 294 294 294 294 294 338 433 497 571 611 611 571 533 433 352 123 137 147 169 194 208 223 256 274 294 315 315 315 315 315 362 464 533 611 655 655 611 571 464 377 124 147 158 181 208 223 239 274 294 315 338 338 338 338 338 388 497 571 655 702 702 655 611 497 404 125 158 169 194 223 239 256 294 315 338 362 362 362 362 362 416 533 611 702 752 752 702 655 533 433 Page 109 ### Page 110 SPL 50 63 80 100 125 160 200 250 315 400 500 630 800 1000 1250 1600 2000 2500 3150 4000 5000 6300 8000 10000 126 169 181 208 239 256 274 315 338 362 388 388 388 388 388 446 571 655 752 806 806 752 702 571 464 127 181 194 223 256 274 294 338 362 388 416 416 416 416 416 478 611 702 806 863 863 806 752 611 497 128 194 208 239 274 294 315 362 388 416 446 446 446 446 446 512 655 752 863 925 925 863 806 655 533 129 208 223 256 294 315 338 388 416 446 478 478 478 478 478 549 702 806 925 991 991 925 863 702 571 130 223 239 274 315 338 362 416 446 478 512 512 512 512 512 588 752 863 991 1062 1062 991 925 752 611 131 239 256 294 338 362 388 446 478 512 549 549 549 549 549 630 806 925 1062 1137 1137 1062 991 806 655 132 256 274 315 362 388 416 478 512 549 588 588 588 588 588 676 863 991 1137 1219 1219 1137 1062 863 702 133 274 294 338 388 416 446 512 549 588 630 630 630 630 630 724 925 1062 1219 1306 1306 1219 1137 925 752 134 294 315 362 416 446 478 549 588 630 676 676 676 676 676 776 991 1137 1306 1399 1399 1306 1219 991 806 135 315 338 388 446 478 512 588 630 676 724 724 724 724 724 832 1062 1219 1399 1499 1499 1399 1306 1062 863 136 338 362 416 478 512 549 630 676 724 776 776 776 776 776 891 1137 1306 1499 1606 1606 1499 1399 1137 925 137 362 388 446 512 549 588 676 724 776 832 832 832 832 832 955 1219 1399 1606 1721 1721 1606 1499 1219 991 138 388 416 478 549 588 630 724 776 832 891 891 891 891 891 1024 1306 1499 1721 1844 1844 1721 1606 1306 1062 139 416 446 512 588 630 676 776 832 891 955 955 955 955 955 1098 1399 1606 1844 1975 1975 1844 1721 1399 1137 140 446 478 549 630 676 724 832 891 955 1024 1024 1024 1024 1024 1176 1499 1721 1975 1975 1844 1499 1219 141 478 512 588 676 724 776 891 955 1024 1098 1098 1098 1098 1098 1261 1606 1844 1975 1606 1306 142 512 549 630 724 776 832 955 1024 1098 1176 1176 1176 1176 1176 1351 1721 1975 1721 1399 143 549 588 676 776 832 891 1024 1098 1176 1261 1261 1261 1261 1261 1448 1844 1844 1499 144 588 630 724 832 891 955 1098 1176 1261 1351 1351 1351 1351 1351 1552 1975 1975 1606 145 630 676 776 891 955 1024 1176 1261 1351 1448 1448 1448 1448 1448 1664 1721 146 676 724 832 955 1024 1098 1261 1351 1448 1552 1552 1552 1552 1552 1783 1844 147 724 776 891 1024 1098 1176 1351 1448 1552 1664 1664 1664 1664 1664 1911 1975 148 776 832 955 1098 1176 1261 1448 1552 1664 1783 1783 1783 1783 1783 2048 149 832 891 1024 1176 1261 1351 1552 1664 1783 1911 1911 1911 1911 1911 150 891 955 1098 1261 1351 1448 1664 1783 1911 2048 2048 2048 2048 2048 Page 110 ### Page 111 Table 10. Example of tone correction calculation for a turbofan engine ### Page 112 1 10 100 1 000 10 000 Total perceived noisiness, , noys N Perceived noise level, PNL, PNdB 150 140 130 120 110 100 90 80 70 60 50 40 Figure 7. Perceived noise level as a function of total perceived noisiness Page 112 ### Page 113 Chapter 4 GUIDELINES FOR TRANSPORT CATEGORY LARGE AIRPLANES, SUBSONIC JETAIRPLANES, AND HELICOPTERS EVALUATED UNDER APPENDIX A OR H OF 14 CFR PART 36 **CHAPTER TABLE OF CONTENTS** 401 EXPLANATORY INFORMATION … 116 401(a) Noise Measurements … 116 401(b) Noise Certification Test and Measurement Conditions … 116 401(b)(1) Ambient temperature … 116 401(b)(2) Calculation of sound attenuation coefficients for the effects of atmospheric absorption … 117 401(b)(3) Wind speed … 120 401(b)(4) Anomalous meteorological conditions … 121 401(b)(5) Time of meteorological measurements … 121 401(b)(6) Aircraft position measurement … 122 401(b)(7) Time space position information (TSPI) measurement system characteristics … 122 401(b)(8) Aircraft performance … 123 401(c) Measurement of Aircraft Noise Received on the Ground … 124 401(c)(1) Environmental specifications … 124 401(c)(2) Measurement system specification … 124 401(c)(3) Microphone orientation … 124 401(c)(4) Microphone specification … 124 401(c)(5) Recorder specifications … 125 401(c)(6) Pre-emphasis … 127 401(c)(7) Attenuator specifications … 127 401(c)(8) Linear integrating analyzer specifications … 127 401(c)(9) Analyzer performance specifications … 127 401(c)(10) Microphone incidence adjustments … 128 401(c)(11) Pink noise specification … 128 401(c)(12) Measurement system field calibration … 128 401(c)(13) Field acoustical calibrations … 128 401(c)(14) Windscreen loss adjustment … 129 401(c)(15) Measurement system background noise … 129 401(d) Calculation of EPNL from Measured Data … 129 401(d)(1) Instantaneous sound pressure levels … 129 401(d)(2) Tone correction calculation … 130 401(d)(3) Adjustments relating to background noise … 130 401(d)(4) Data resolution after tone correction calculation … 130 401(d)(5) Band sharing adjustment … 130 401(d)(6) Calculation of band sharing adjustment … 131 401(d)(7) Noise Duration … 131 401(d)(8) Equations for computing duration correction … 132 401(e) Reporting of Data to the FAA … 133 401(e)(1) Compliance records … 133 401(e)(2) EPNL average values … 133 401(e)(3) Calculation of 90 percent confidence intervals … 134 401(f) Nomenclature: Symbols and Units … 134 401(g) Sound Attenuation in Air … 135 401(h) Adjustment of Helicopter Flight Test Results … 135 401(h)(1) Adjustments to reference conditions … 135 401(h)(2) Reference data sources … 136 401(h)(3) Adjustments to measured noise data … 136 401(h)(4) Takeoff profile … 136 ### Page 114 401(h)(5) Takeoff test conditions … 137 401(h)(6) Takeoff flight test procedures … 139 401(h)(7) Flyover configuration … 142 401(h)(8) Flyover test procedure … 144 401(h)(9) Approach configuration … 148 401(h)(10) Approach test condition … 149 401(h)(11) Approach test procedure … 150 401(h)(12) Adjustments of PNL and PNLT … 152 401(h)(13) Duration adjustment to EPNL … 152 401(i) Adjustment of Airplane Flight Test Results … 153 401(i)(1) Adjustments to reference conditions … 153 401(i)(1)(A) Origin of reference data … 154 401(i)(2) Takeoff configuration/profile … 154 401(i)(3) Takeoff test procedure … 157 401(i)(4) Approach test configuration … 157 401(i)(5) Approach test procedure … 159 401(i)(6) Lateral noise measurement … 160 401(i)(7) Integrated procedure adjustment … 160 401(i)(8) Acoustic emission angles … 160 402 EQUIVALENT PROCEDURES INFORMATION … 160 402(a) Subsonic Jet Airplanes … 160 402(a)(1) Flight test procedures … 161 402(a)(1)(A) Flight path intercepts … 161 402(a)(1)(A)(i) For takeoff … 161 402(a)(1)(A)(ii) For approach … 161 402(a)(1)(B) Generalized flight test procedures … 162 402(a)(1)(B)(i) For derivation of noise-power-distance (NPD) data … 162 402(a)(1)(B)(ii) For flight test procedures for determination of changes … 164 402(a)(1)(C) Determination of the lateral certification noise levels … 164 402(a)(1)(D) Takeoff flyover noise levels with thrust (power) reduction … 165 402(a)(1)(E) Measurements at non-reference points … 166 402(a)(1)(F) Atmospheric test conditions … 166 402(a)(1)(G) Layering equivalency … 166 402(a)(2) Analytical procedures … 167 402(a)(2)(A) Flyover noise levels with thrust (power) reduction … 167 402(a)(2)(B) Equivalent procedures based upon analytical methods … 168 402(a)(2)(C) Equivalent procedure for calculating the certification noise levels … 169 402(a)(3) Static engine noise tests and projections to flight noise levels … 170 402(a)(3)(A) General … 170 402(a)(3)(B) Limitation on the projection of static-to-flight data … 170 402(a)(3)(C) Static engine noise test procedures … 172 402(a)(3)(C)(i) General … 172 402(a)(3)(C)(ii) Test site characteristics … 172 402(a)(3)(C)(iii) Data acquisition and reduction systems … 173 402(a)(3)(C)(iv) Microphone types, locations and installations … 174 402(a)(3)(D) Projection of static engine data to airplane flight conditions … 174 402(a)(3)(D)(i) General … 174 402(a)(3)(D)(ii) Normalization to reference conditions … 176 402(a)(3)(D)(iii) Separation into broadband and tone noise … 177 402(a)(3)(D)(iv) Separation into contributing noise sources … 177 402(a)(3)(D)(v) Noise source position effects … 178 402(a)(3)(D)(vi) Engine flight conditions … 179 402(a)(3)(d)(vii) Noise source motion effects … 179 402(a)(3)(D)(viii) Airplane configuration effects … 180 ### Page 115 402(a)(3)(D)(ix) Airframe noise … 181 402(a)(3)(D)(x) Airplane flight path considerations … 181 402(a)(3)(D)(xi) Total noise spectra … 181 402(a)(3)(D)(xii) EPNL computations … 182 402(a)(3)(D)(xiii) Changes to noise levels … 182 402(b) Propeller-Driven Airplanes Over 19,000 lb … 182 402(b)(1) Flight test procedures … 182 402(b)(1)(A) Flight path intercept procedures … 182 402(b)(1)(B) Generalized flight test procedures … 182 402(b)(1)(C) Determination of the lateral certification noise level … 183 402(b)(1)(D) Measurements at non-reference points … 184 402(b)(2) Analytical procedures … 185 402(b)(3) Ground static testing procedures … 186 402(b)(3)(A) General … 186 402(b)(3)(B) Guidance on the test-site characteristics … 186 402(b)(3)(C) Static tests of the gas generator … 186 402(c) Helicopters … 186 402(c)(1) Flight test procedures … 187 402(c)(1)(A) Helicopter test speed … 187 402(c)(1)(B) Atmospheric test conditions … 187 402(c)(1)(C) Temperature and relative humidity measurements … 187 402(c)(1)(D) Modifications or upgrades involving aerodynamic drag changes … 188 402(c)(1)(E) Anomalous test conditions … 188 402(c)(1)(F) Helicopter test rotor speed … 189 402(c)(1)(G) Helicopter test weight … 189 402(c)(1)(H) Helicopter approach … 189 402(c)(2) Analytical procedures … 189 402(c)(2)(A) Helicopter test window for zero adjustment for atmospheric attenuation … 189 402(c)(2)(B) Procedure for the determination of source noise adjustment … 191 403 TECHNICAL PROCEDURES INFORMATION … 192 403(a) Jet and Propeller-Driven Airplanes and Helicopters … 192 403(a)(1) Adjustment of measured noise data to reference conditions … 192 403(a)(2) Determination of reference-condition noise geometry … 192 403(a)(2)(A) Assumptions … 193 403(a)(2)(B) Steps involved … 193 403(a)(2)(C) Characterization of a straight-line average flight path based on descriptors … 193 403(a)(2)(D) Determination of test aircraft position at time of sound emission … 195 403(a)(2)(E) Calculation of the geometric minimum distance between the test flight path … 197 403(a)(2)(F) Determination of test aircraft noise geometry … 200 403(a)(2)(G) Determination of reference flight path … 201 403(a)(2)(H) Determination of reference sound propagation distance … 202 403(a)(2)(I) Determination of effective duration for each acoustic spectrum … 203 403(a)(3) Computation of EPNL by the simplified method of adjustment … 205 403(a)(3)(A) Inputs required for performing the simplified method … 205 403(a)(3)(B) Adjustment of sound pressure levels to reference conditions … 206 403(a)(3)(C) Determination of reference-condition tone-corrected perceived noise levels … 206 403(a)(3)(D) Application of test-day bandsharing adjustment to reference-condition … 206 403(a)(3)(E) Determination of the ΔPeak simplified adjustment term to account for the effects … 207 403(a)(3)(G) Determination of the Δ3 simplified adjustment term to account for differences … 207 403(a)(3)(H) Determination of the simplified reference-condition effective perceived noise level … 208 403(a)(3)(I) Worked example calculations illustrating the simplified method … 208 403(a)(4) Computation of EPNL by the integrated method of adjustment … 208 403(a)(4)(A) Inputs required for performing the integrated method … 208 403(a)(4)(B) Adjustment of sound pressure levels to reference conditions … 209 ### Page 116 403(a)(4)(C) Determination of reference-condition tone-corrected perceived noise … 209 403(a)(4)(D) Identification of integrated reference-condition maximum tone-corrected … 209 403(a)(4)(E) Identification of integrated reference-condition EPNL Noise Duration … 210 403(a)(4)(F) Summation of integrated reference-condition EPNL … 210 403(a)(4)(G) Worked example calculations illustrating the integrated method … 210 403(b) Jet Airplanes … 211 403(b)(1) Control of noise certification computer program software and documentation … 211 403(b)(1)(A) General … 211 403(b)(1)(B) Software control procedures — four key elements … 212 403(b)(1)(B)(i) Configuration index … 212 403(b)(1)(B)(ii) Software control plan … 212 403(b)(1)(B)(iii) Design description … 212 403(b)(1)(B)(iv) Verification process … 212 403(b)(1)(C) Applicability … 212 403(b)(2) Identification of spectral irregularities … 212 403(b)(2)(A) Introduction … 213 403(b)(2)(B) Methods for identifying false tones … 213 403(b)(2)(B)(i) Frequency tracking … 213 403(b)(2)(B)(ii) Narrow-band analysis … 214 403(b)(2)(B)(iii) Microphone mounting height … 214 403(b)(2)(B)(iv) Inspection of noise time-histories … 214 403(b)(2)(C) Treatment of false tones … 214 403(b)(3) Noise data adjustments for tests at high altitude test sites … 214 403(b)(3)(A) Introduction … 214 403(b)(3)(B) Jet source noise adjustment … 214 403(b)(3)(B)(i) Criteria … 215 403(b)(3)(B)(ii) Adjustment procedures … 215 EXPLANATORY INFORMATION [ETM 4.1] Noise Measurements 401(a) §A36.2 [ETM 4.1.1] See 301 for technical procedures generally applicable for noise certification tests of all aircraft types including those evaluated under the provisions of Appendices A or H of part 36. Procedures specific to Appendix A or H are presented in the following section. Noise Certification Test and Measurement Conditions 401(b) §A36.2 [ETM 4.1.2] 401(b)(1) Ambient temperature GM §A36.2.2.2(b) [ETM GM A2 2.2.2.4.1 b)] (1) Criteria for measuring atmospheric conditions Experience has shown that proper measurement of non-r eference meteorological conditions and the associated adjustment of noise data for these conditions are crucial to obtaining accurate, consistent and repeatable test results. For airplanes, meteorological observations of the temperature and relative humidity are required over the whole sound propagation path from the aircraft to the vicinity of the noise measurem ent points. For helicopters, temperature and relative humidity measurements are require d at 33 ft (10 m) in the vicinity of the noise measurement points. AMC §A36.2.2.2(b) [ETM AMC A2 2.2.2.4.1 b)] Page 116 ### Page 117 (1) Atmospheric measurements Several methods have been approved for the measurement of atmospheric conditions from 33 ft (10 m) above the ground to the altitude of the test airplane. Some appli cants have used instrument ed balloons. Another method consists of a meteorological airplane, manned or unmanned, flown in a spiral flight path in the vicinity of the noise measurement points to measure the dry bulb temperature and dew point along the sound propagation path. 401(b)(2) Calculation of sound attenuation coefficients for the ef fects of atmospheric absorption AMC §A36.2.2.2(d) [ETM AMC A2 2.2.2.2.1)] (1) Basic data Measurements of the ambient temperature and relative humidity should be made at 33 ft (10 m) above the ground. The ambient temperature and relative humidity should also be determined with vertical height increments not greater than 100 ft (30 m) over the sound propagation path. All measurements of ambient temperature and relative humidity must be obtained within 30 minutes of each airplane test run. (2) Determination of the average sound attenuation coefficient Table 11 is an example of calculation of sound attenuation coefficients in the 3150 Hz one-third octave band for an airplane approach noise certification when multiple layering is not required. Temperature and humidity values obtained from atmospheric soundings performed before and after a series of airplane test runs are interpolated to the time of PNLTM. Table 11. Basic data (layering not required) Height Temperature Relative humidity α (3150 Hz) (m) (°C) (%) (dB/100 m) 10 14.1 50 2.45 30 13.4 53 2.38 60 12.9 56 2.30 90 12.2 57 2.33 120 11.5 58 2.37 150 11.3 61 2.27 The individual coefficients shown in Table 11are calculated at vertical height increments of 30 m from 10 m to a height of 150 m. The ambient conditions from the ground to 10 m are assumed to be those measured at 10 m. The individual sound attenuation coefficients for the 3150 Hz one-third octave band shown in Table 11 vary by less than 0.5 dB/100 m relative to the value determined at 10 m (33 ft). In this case the coefficient to be used for adjustment of sound pressure levels fr om test to reference conditions is the average of the coefficients at 10 m (33 ft) and at the height of the airplane at the time of PNLTM. For this example, the height of the test airplane at the time of PNLTM is 125 m. The associated attenuation Page 117 ### Page 118 coefficient is calculated by linear interpolation as follows: ݕൌ ݕ௔ ൅ ሺݔെݔ ௔ሻ ∗ሺ ݕ௕ െݕ ௔ሻሺݔ௕ െݔ ௔ሻ ߙሺ3150ሻଵଶହ௠ ൌ2 . 3 7 ൅ ሺ125 െ 120ሻ ∗ ሺ2.27 െ 2.37ሻሺ150 െ 120ሻ ߙሺ3150ሻଵଶହ௠ ൌ 2.35/ܤ݀100݉ . Then the average attenuation coefficient for the 3150 Hz one-third octave band used for adjustment of the airplane sound pressure levels is calculated as follows: ߙሺ3150ሻ ൌ ߙሺ3150ሻଵ଴௠ ൅ ߙሺ3150ሻଵଶହ௠ 2 ߙሺ3150ሻ ൌ 2.45 ൅ 2.352 ߙሺ3150ሻ ൌ 2.40/ܤ݀100݉ . The coefficients for the other one-third octave band s are determined in a similar manner. These average coefficients are then used in the adjustment of airplane SPLs to reference conditions. The same general procedure would be used if no layering was required for determin ing the average coefficients during flyover and lateral noise certification measurements. (3) Determination of the cumulative sound attenuation coefficients Table 12 is an example of calculation of sound attenuation coefficients in the 3150 Hz one-third octave band for an airplane flyover noise certification when multiple layering is required. Temperature and humidity values obtained from atmospheric soundings performed before and after a series of airplane test runs are interpolated to the time of PNLTM. Table 12. Basic data (layering required) Height Temperature Relative humidity α (3 150 Hz) (m) (°C) (%) (dB/100 m) 10 7.2 80 2.09 30 7.2 75 2.23 60 8.9 73 2.11 90 10.0 67 2.19 120 10.6 63 2.27 150 10.6 62 2.31 180 10.6 61 2.34 210 10.6 59 2.43 240 11.1 55 2.57 270 11.7 53 2.59 Page 118 ### Page 119 Height (m) 300 330 360 390 420 Temperature (°C) 11.7 11.1 11.1 11.1 11.1 Relative humidity (%) 51 51 50 47 46 α (3 150 Hz) (dB/100 m) 2.70 2.79 2.84 3.04 3.10 The individual coefficients shown in Table 12 are calculated at vertical height increments of 30 m from 10 m to a height of 420 m. The ambient conditions from the ground to 10 m are assumed to be those measured at 10 m. The individual sound attenuation coefficients for the 3150 Hz one-third octave band shown in Table 12 vary by more than 0.5 dB/100 m. In this case the coefficient to be used for adjustment of sound pressure levels from test to reference conditions is the cumulative sound attenuation from the ground to the height of the airplane at the time of PNLTM. In the absence of extreme or anomalous conditions (e.g., large variations in, or inversions of, temperature and/or humidity), which will generally be the case, it is acceptable, subject to the approval by the FAA, to determine the cumulative sound attenuation coefficients for each one-third octave band from a simple average of the coefficients at the boundaries of each layer. Where extreme or anomalous conditions are present (e.g., large variations in, or inversions of, temperature and/or humidity) the cumulative sound attenuation coefficients for each one-third octave band should be determined by apportioning the sound attenuation coefficients for each layer. Table 13 illustrates an example of such a method. Table 13. Determination of the cumulative sound attenuation Average layer Apportioned Effective sound sound Effective layer Sound attenuation attenuation attenuation layer depth coefficients, coefficients, coefficients, Layer depth proportion α (3 150 Hz) α (3 150 Hz) α (3 150 Hz) boundaries (m) (%) (dB/100 m) (dB/100 m) (dB/100 m) 0–30 28.8 7.03 2.09–2.23 2.16 0.1518 30–60 30.0 7.32 2.23–2.11 2.17 0.1589 60–90 30.0 7.32 2.11–2.19 2.15 0.1574 90–120 30.0 7.32 2.19–2.27 2.23 0.1633 120–150 30.0 7.32 2.27–2.31 2.29 0.1676 150–180 30.0 7.32 2.31–2.34 2.32 0.1698 180–210 30.0 7.32 2.34–2.43 2.39 0.1750 210–240 30.0 7.32 2.43–2.57 2.50 0.1830 240–270 30.0 7.32 2.57–2.59 2.58 0.1889 270–300 30.0 7.32 2.59–2.70 2.65 0.1940 300–330 30.0 7.32 2.70–2.79 2.74 0.2006 330–360 30.0 7.32 2.79–2.84 2.82 0.2064 Page 119 ### Page 120 Average layer Apportioned Effective sound sound Effective layer Sound attenuation attenuation attenuation layer depth coefficients, coefficients, coefficients, Layer depth proportion α (3 150 Hz) α (3 150 Hz) α (3 150 Hz) boundaries (m) (%) (dB/100 m) (dB/100 m) (dB/100 m) 360–390 30.0 7.32 2.84–3.04 2.94 0.2152 390–411 21.0 5.12 3.04–3.08 3.06 0.1568 Cumulative 2.49 sound attenuation coefficient, α (3150 Hz) (dB/100 m): The atmosphere is first divided into layers from the ground to the airplane height. For this example the height of the airplane at the time of PNLTM is 411 m. The sound attenuation coefficient at the height of the test airplane is calculated by linear interpolation of the sound attenuation coefficients at the upper and lower boundaries of the uppermost layer. The effective layer depth is determined as follows: for all layers between the airplane and the microphone, except the lowest layer containing the microphone and the upperm ost layer containing the airplane, the effective layer depth is the full 30 m; for the lowest layer containing the microphone, the effective layer depth is 30 m minus the 1.2 m height of the microphone; for the uppermost layer containing the airplane, the effective layer depth is the height of the airplane minus the height of the lower boundary of the layer. The effective layer depth proportion for each layer is determined as the ratio of that layer’s effective depth relative to the total vertical component of the sound propagation distance from the microphone to the height of the airplane at the time of PNLTM. The average sound attenuation coefficient for each layer is obtained by averaging the coefficients at the upper and lower boundaries of the layer. The apportioned sound attenuation coefficient for each layer is obtained by multiplying the average layer sound attenuation coefficient by the effective layer depth proportion. The summation of all apportioned sound attenuation coefficients results in the cumulative sound attenuation coefficient. In this example the cumulative coefficient is calculated for the 3 150 Hz one-third octave band. The same general procedure would be used to obtain the cumulative sound attenuation coefficient for each one-third octave band. These coefficients are then used in the adjustment of airplane SPLs to reference conditions. 401(b)(3) Wind speed AMC §A36.2.2.2(e) [ETM AMC A2 2.2.2.4.1 e) f) g) h)] (1) Real-time crosswind component measurements Applicants are advised to provide ap proved real-time crosswind component measurement systems such that the crosswind component speeds can be verified after each aircraft test run. When the applicant uses a wind measurement system that is remotely located and not readily accessible, such as chart recorders that simultaneously and independently measure and record wind speed and direction, it may not be practical to determine the real-time crosswind component for each test run. For airplanes, if the applicant does not provide an acceptable real-time crosswind component measurement system, the 10 kt (5.1 m/s) maximum crosswind component and the 7 kt (3.6 m/s) average crosswind component become the maximum wind limitations regardless of wind direction. ### Page 121 401(b)(4) Anomalous meteorological conditions GM §A36.2.2.2(f) [ETM GM A2 2.2.2.4.1 i)] (1) Anomalous winds For airplanes, compliance of measured wind speeds with the requirements of A36.2.2.2 e) of part 36 may not be sufficient to ensure that the wind speeds at the airplane height or along the sound propagation path are not excessive. Such conditions may exist as a steady headwind, tailwind or crosswind, or as a wind from varying directions with increasing height. Anomalous winds may affect the handling characteristics of an aircraft during the noise duration. They also may affect the transmitted noise. Anomalous winds include not only gusts and turbulent winds, but also wind shear, strong vertical winds, and high crosswinds at the aircraft height and along the sound propagation path. An applicant may be required to measure winds aloft and provide FAA with the information. Acceptability of the wind conditions over the propagation path will be determined by the (see A36. 2.2.2(f) of part 36). (2) Winds aloft measurement Modern INS and DGPS can provide on-board aircraft data that can be used to quantify winds aloft. The measurement of winds aloft can further be processed to provide a permanent record of wind speed and direction. (3) Effects of wind on airplane control FAA has permitted a ±20 percent tolerance in overhead test height and a ±10° lateral tolerance relative to the extended runway centerline. If the flight crew cannot fly within the pre-test-approved flight path tolerance limits, or experiences major variations in airspeed, or the airplane crabs or yaws significantly during the flight, adverse or anomalous wind conditions aloft are often the cause. (4) Effects of wind on helicopter control If the test helicopter cannot be flown within the pre-test -approved flight path tolerance limits, or experiences major variations in airspeed, or the aircraft yaws or sideslips excessively during the flight, adverse or anomalous wind conditions aloft are often the cause. Normally such issues arise only with gusty wind conditions, high crosswinds or in the presence of strong thermals. AMC §A36.2.2.2(f) [ETM AMC A2 2.2.2.4 i)] (1) Flight path The flight crew should observe and record any occurrence where conditions aloft cause difficulty in maintaining the flight path or airspeeds, or when rough air in general makes the flight unacceptable. In the context of determining whether such conditions are present for airplanes, B36.8(g) of part 36 specifies that “for take-off, lateral, and approach conditions, the vari ation in instantaneous IAS of the airplane must be maintained within ±3 percent of the average airspeed between the 10 dB-down points. This shall be determined by reference to the pilot’s airspeed indicator. However, when the instantaneous IAS varies from the average airspeed over the 10 dB-down points by more than ±3 kt (±5.5 km/h), and this is judged by the FAA representative on the flight deck to be due to atmospheric turbulence, then the flight so affected shall be rejected for noise certification purposes.” 401(b)(5) Time of meteorological measurements GM §A36.2.2.2(g) [ETM GM A2 2.2.2.2.1] (1) Upper atmospheric condition measurements Page 121 ### Page 122 Atmospheric conditions affect sound propagation. Therefore, measurements of temperature and relative humidity must be made before and after each aircraft test run, at least one of which must be made within 30 minutes of the test run. To avoid the possibility that the meteorological conditions might change significantly over time, both measurements must be representative of the prevailing conditions during the test run. The measurements must be made using an approved method at 33 ft (10 m) above the ground surface and, for airplanes only, from 33 ft (10 m) above the ground surface to the airplane test he ight at time of PNLTM. These measurements must be obtained and validated through out the test period to en sure acceptable meteorological data for the noise data evaluation process. AMC §A36.2.2.2(g) [ETM AMC A2 2.2.2.2.1] (1) Atmospheric measurements Applicants should consider the maximum height that will be attained within the next 60 minutes, or less, of airplane test runs to ensure that adequate upper at mospheric measurements are acq uired. Interpolations of atmospheric data for all test runs are made to the airp lane height at the time of PNLTM. To have sufficient meteorological data to perform the interpolation to the ac tual time of each test run, the first meteorological measurement flight of the day should be made not earlier than 30 minutes before the first test run, and the last meteorological measurement flight of the day should be made not later than 30 minutes after the last test run flight of the day. (2) Atmospheric data interpolation The temperature and relative humidity data at the actual time of the test run must be interpolated over time and height, as necessary, from the measured meteorological data. The interpolation time of the test run may be taken to be either the time the aircraft flew overhead or abeam the noise measurement point, or the time of PNLTM. 401(b)(6) Aircraft position measurement GM §A36.2.3.1 [ETM GM A2 2.3.1] (1) Independent aircraft position determination The FAA will approve only those aircraft position and height indicating and recording systems that are independent from the direct aircraft flight path indicating systems. The data from such independent systems should be recorded to produce a time-coordinated permanent record of each test. The independent system restriction does not prohibit use of real-time flight guidance systems (e.g., CDI or GDI) on-board the aircraft to assist the flight crew during noise certification tests. Systems such as microwave space position systems, INS, precision distance measuring uni t (DMU), and DGPS can also provide guidance to the flight crew by providing the direct, real-time aircraft position relative to the extended runway centerline. 401(b)(7) Time space position information (TSPI) measurement system characteristics GM §A36.2.3.2 [ETM GM A2 2.3.2] (1) Measurement system synchronization Approved aircraft position and height measurement systems must be time-synchronized with the noise and meteorological measurement systems. The time synchronization between noise measurements and aircraft position should be precise. A common time base should be used to synchronize noise, aircraft tracking and meteorological measurements (see §A36.2.3.2(b)(1)(3)(3)) of part 36 for details). TSPI should be determined at intervals no greater than one-half second throughout the sound-measuring period (i.e., within 10 dB of PNLTM) by an approved method that is independent from systems installed aboard, and Page 122 ### Page 123 normally used to control, the aircraft. During processing, measured TSPI data must be interpolated over time to the time of sound emission of each one-half second noise data record within the 10 dB-down period. The time associated with each one-half second record is 0.75 seconds before the end of each 2-second exponential averaging period (see A36.3.7.6 of part 36). Although the simplified procedure requir es adjustment of only the (PNLT) maximum record to the reference track, sound emission coordinates should be determined for each one-half second record for use in background noise adjustment procedures and/or for determination of incidence-dependent free-field microphone and windscreen adjustments. (2) Measurement system component approval Some off-the-shelf TSPI equipment may require software enhancement to accommodate the specific installation. Each applicant should submit information to the FAA about the software used. The FAA will determine whether the software yields results that satisfy the part 36 Standards. All TSPI equipment and software should be demonstrated to, and approved by, the FAA to ensure the system’s operational accuracy. (3) Methods of time synchronization Special care should be taken to properly synchronize noise data recordings with TSPI data (see 304 of this AC for details of specific methods). 401(b)(8) Aircraft performance GM §A36.2.3.3 [ETM GM A2 2.3.3] (1) Aircraft and engine performance parameters Examples of parameters needed for measurement of aircraft and engine performance include aircraft height, climb angle, airspeed, gross weight, flap position, landing gear position, engine thrust (power) setting parameters (e.g., compressor rotor speed, engine pressure ratio, exhaust ga s temperature), and aircraft accessory condition (e.g., air conditioning and auxiliary power unit (APU) “on” or “off”). Any other parameters that may affect measurement or adjustment of noise data and/or aircraft or engine performance should also be recorded throughout the 10 dB-down period (e.g., the status of surge bleed valves (SBV) and the center of gravity (CG) position). AMC §A36.2.3.3 [ETM AMC A2 2.3.3] (1) Aircraft performance measurements Calibrated instrumentation is required to determine aircraft performance. Adequate aircraft and engine parameters are to be recorded during all certification testing to ensure that aircraft performance can be accurately determined. For example, for transport airplanes this may necessitate measurement and recording of flap position, landing gear position, speed brake position, APU operation, and normal engine thrust (power) setting and associated flight parameters. Determination and recording of adequate information enables validation of the test configuration and adjustment of performance and engine performance from test conditions to reference conditions specified in B36.7(a)(5) of part 36. (2) Recorder sampling rate The measurements of aircraft position, airspeed, perf ormance and engine performance parameters are to be recorded at an approved sampling rate sufficient to permit adjustments from test to reference conditions throughout the 10 dB-down period. An acceptable recording sampling rate for transport category airplanes is two to five samples per second. ### Page 124 401(c) Measurement of Aircraft Noise Received on the Ground §A36.3 [ETM 4.1.3] 401(c)(1) Environmental specifications GM §A36.3.2.1 [ETM GM A2 3.2] (1) Measurement system performance The environmental conditions for specifying the performance of a measurement system are specified in A36.3.2.1 of part 36. 401(c)(2) Measurement sy stem specification GM §A36.3.3.1 [ETM GM A2 3.3.1] (1) Measurement system criteria The specifications for a measurement system allow flexibility in the procurement of measurement system components by the applicant. While on-site EPNL analysis may be useful for estimation of recording levels or for other diagnostic purposes, a true acoustical analysis requires that data be recorded in the field. This will allow for later re-analysis or auditing of acoustical data. A recording also facilitates later offline processing of acoustic data, including application of adjustments for items such as system frequency response, microphone pressure response and analyzer bandwidth error. Recording simplifies synchronization with other pertinent data, such as tracking and meteorological measurements. Such synchronization is necessary for proper application of many of the required adjustments to noise data, such as adjustments for microphone free-field response, windscreen incidence-dependent insertion loss, the influence of ambient noise, high altitude jet noise effects, non-reference flight performance, and non-reference meteorological conditions. (2) Approval of measurement system The approval by the FAA should be obtained for systems used for measurement, recording and analysis of aircraft noise. Most of the currently available system components that are appropriate for aircraft noise certification have already been approved, but implementation of new technology and variants or upgrades of existing components may require approval by the FAA. Of special concern is the potential for a digital component’s functionality to change as a result of firmware or operating system upgrades or modifications. Applicants should be aware that approval of a particular component might be version-dependent. 401(c)(3) Microphone orientation AMC §A36.3.5.2 [ETM AMC A2 3.5.2] (1) Microphone orientation Figure 8 shows the orientations relative to a microphone sensing unit for grazing and normal incidence. For microphones located directly under the flight path, an orientation angle of 90° from vertical is appropriate regardless of target height. For noise measurements to the side of the flight path, applicants may wish to reorient the microphones for grazing incidence for each target height in order to maintain substantially grazing incidence throughout the 10 dB-down periods. In many cases, this reorientation can eliminate the need to apply data adjustments for varying incidence, since the incidence angles will be more likely to be contained within ±30º of grazing incidence. Figure 8 provides illustrations of microphones positioned for grazing incidence under the flight path and to the side of the flight path of an airplane. 401(c)(4) Microphone specification GM §A36.3.5.4(b) [ETM GM A2 3.5.4] (1) Microphone specifications Page 124 ### Page 125 Table A36-1 specifies the maximum permitted differences between the free-field sensitivity of a microphone at normal incidence and the free-field sensitivity at specified sound incidence angles for sinusoidal sound waves at each one-third octave band nominal midband frequency over the range of 50 Hz to 10 kHz. These differences are larger at higher frequencies, allowing for the effect of the microphone body in a free-field environment. (2) Microphone characteristics The specifications of Table A36-1 are based on the performance characteristics of typical one-half inch condenser microphones designed for nearly uniform frequency response at grazing incidence (see Figure 8). Other microphones may be used, provided they meet the specified performance requirements. For example, prepolarized (i.e., electret condenser) free -field microphones greatly minimize the possibility of arcing in humid environments and do not require an external polarization voltage. Although many of these microphones are intended primarily for use in normal-incidence free-fiel d applications, they can be used in aircraft noise certification testing if their performance at grazing incidence meets the requirements of A36.3.5 of part 36. 401(c)(5) Recorder specifications GM §A36.3.6.1 [ETM GM A2 3.6.1] (1) Recorder types An applicant has a choice of recorder types that will satisfy the requirement for recording “the complete acoustic signal” during certification testing. In addition to a magnetic tape recorder, other means of attaining a “true” acoustic recording include digital audiotape (DAT), reco rdable compact disc (CD-R) and direct-to-hard-disk recording. The applicant should be aware that systems that use data compression techniques that result in substantial data loss, such as mini-disc (MD) or digital compact cassette (DCC), are not acceptable. AMC §A36.3.6.1 [ETM AMC A2 3.6.1] (1) Frequency range for recordings The time-varying waveform produced by the microphone response to noise signals during certification tests should be recorded. If there are questions about the data observed during the tests, the recording can be replayed, multiple times if necessary, to verify the results. Recorded data, whether digital or analog in nature, should allow reproduction and reprocessing of an analog signal over the frequency range of 40 Hz to 12.6 kHz. A dynamic range of at least 60 dB is recommended. Many typical instrumentation DAT recorders feature a nominal 10 kHz bandwidth operating mode in which the attenuating response of the anti-aliasing filter intrudes within the 10 kHz one-third octave passband. In such cases, the recorder should be operated in a nominal 20 kHz bandwidth mode, which may reduce the number of available channels or the duration of available time per tape. Note.— Although the one-third octave bands of interest are those with nominal center frequencies of 50 Hz through 10 kHz, to ensure that the entire actual band width of the uppermost and lowermost bands is included, the center frequencies of the one-third octave bands immediately outside this range are specified. ### Page 126 0° normal incidence Sensing element 180° 120° 60° 90° grazing incidence Figure 8. Illustration of sound incidence angles on a microphone (2) Digital recording levels The overload characteristic of a digita l system is determined primarily by the limits of the analog-to-digital conversion. Since such an overload co ndition is characterized by an abrupt, catastrophic type of distortion, the level range should be set so that the anticipated maximu m signal level is at least 10 dB, and preferably 20 dB, below the upper boundary of the linear operating range. (3) Dynamic range limits for digital recorders The lower limit of a digital recording system’s usable dynamic range is more often determined by amplitude nonlinearity due to “quantization error”, rather than by the presence of a noise floor. Digital devices such as recorders or analyzers that are to be used for aircraft noise certif ication purposes should be tested to determine the extent of such non-linearity. (4) 16-bit quantization systems The theoretical dynamic range of such a system is usually assumed to be near 96 dB (i.e., 20 log (2 16)). At the lower limit of this range, there is a potential for a 6 dB error in the digitized signal versus the analog input signal that it represents. Reference T1 imposes a ±0.4 dB limit on acceptable linearity error in the reference level range and ±0.5 dB for a linear operating range of at least 50 dB. As amplitude levels are increased above the lower quantization limit, the linearity error is reduced. If the guidance for setting the level range is followed, the usable Page 126 ### Page 127 dynamic range is further decreased. Significant improv ement of amplitude linearity can be obtained via the implementation of techniques such as oversampling and dithering. Therefore, testing must be performed to determine the actual limits for each digital recording system. Note that assumptions based on experience with analog systems do not always apply. 401(c)(6) Pre-emphasis AMC §A36.3.6.2(a) [ETM AMC A2 3.6.2] (1) Pre-emphasis systems Use of pre-emphasis systems will be allowed only if the system also employs complementary de-emphasis. Attempts to compensate for the effects of a pre-emphasis filter by applying one-third octave band de-emphasis adjustments, either numerically to analyzed data via a pink noise adjustment, or on a band-by-band basis using separate gain stages for each one-third octave band filter, are not allowed. In addition, use of a pre-emphasis/deemphasis system will require testing and documentation of all filters and gain stages involved to ensure that any errors are quantified and minimized and that the system performs predictably and reliably. 401(c)(7) Attenuator specifications AMC §A36.3.6.9 [ETM AMC A2 3.6.9] (1) Attenuator specifications Attenuator specifications allow for the use of switchable voltage input range settings, now commonplace on DAT recorders, as controllable attenuation steps for gain-settin g purposes. In all cases, attenuators should have fixed repeatable steps. Any devices in the measurement system that use vernier or continuously-adjustable gain controls should also have some demonstrable means of being fixed, or locked at a specific setting, to eliminate non-traceable gain errors. 401(c)(8) Linear integrating analyzer specifications GM §A36.3.7.2(c) [ETM GM A2 3.7.2] (1) Externally controlled linear-integrating analysers In cases where a computer or other external device is used to control and/or communicate with an analyser performing linear integration, extra care should be taken to ensure that the integration period requirements are met. Some analysers from major manufacturers have required a factory modification in order to provide an integration time within 5 ms of the specified 500-ms integration period. 401(c)(9) Analyzer perf ormance specifications GM §A36.3.7.4 [ETM GM A2 3.7.3] (1) Analyzer specifications Reference T2 specifies the electrical performance requi rements of one-third octave band filters, including tolerances for the attenuation in the transition bands (i.e ., “skirts”) adjacent to the one-third octave passbands. Most digital one-third octave band analysis systems o ffer only hardwired filtering algorithms that emulate the response of a traditional third-order analysis filter havi ng a maximally flat passband. However, some analysis systems allow the selection of other filtering algorithms which might not provide equivalent performance. Applicants should demonstrate the effects that alternate filter design response characteristics might have on noise certification EPNL values. (2) Determination of bandwidth error adjustments The manufacturer can establish the geometric center frequen cies of one-third octave band filters using either Page 127 ### Page 128 Base 2 or Base 10 systems. While the use of either method results in frequencies close to the nominal center frequencies referred to in Table A36-3, it is important to note which system is used so that the bandwidth error adjustment can be properly determined. Use of test frequencies calculated by a different base-number system than that for which the analyzer was designed can result in erroneous values for these adjustments. 401(c)(10) Microphone incidence adjustments AMC §A36.3.9.3 [ETM AMC A2 3.9.3] (1) Applications of adjustments for incidence When using microphones whose frequency response is nearly flat at grazing incidence, and when the angles of incidence of sound emitted from the aircraft are within ±30 of grazing incidence, a single set of data adjustments for free-field response and windscreen insertion loss, based on grazing incidence, is considered sufficient to account for incidence effects. When it is impractical to orient the micr ophone properly to maintain grazing incidence, provided that a continuous record of TSPI is available, free-field and windscreen insertion-loss incidence data adjustments can be applied to the noise da ta on a spectral-record by spectral-record basis. These adjustments are obtained by calculating the angle of in cidence for each record, usin g the point of time which characterizes the 2-second averaging period (see A36.3.7.6 of part 36) and determining the aircraft’s sound emission coordinates and angle of incidence for the sound measured at that time. 401(c)(11) Pink noise specification GM §A36.3.9.4 [ETM GM A2 3.9.4] (1) Pink noise Pink noise contains equal energy in each octave band or fractional octave band (e.g., the octave from 100 Hz to 200 Hz contains the same amount of energy as the octave from 1 kHz to 2 kHz, although for the lower-frequency octave, it is distributed over a frequency range 10 times narrower). (2) Pink noise usage Because of the dynamic nature of the pink noise signal, longer samples produce statistically better measurements. A minimum duration of 30 seconds of pink noise should be recorded. 401(c)(12) Measurement sy stem field calibration AMC §A36.3.9.5 [ETM AMC A2 3.9.5] (1) Measurement system field calibration (all components of the measurement system, except microphones) All components of the measur ement system, except microphones, should be tested while deployed in the field, using pink noise at a level within 5 dB of the calibration level (see A36.3.9.5 of part 36). The signal should be recorded for a duration of at least 30 seconds so that one-third octave band system frequency response adjustments can be determined and applied during analysis. The pink noise generator should be calibrated within 6 months of the measurement and is acceptable for certification use only if its output in each one-t hird octave band does not change by more than 0.2 dB between calibrations. 401(c)(13) Field acoustical calibrations AMC §A36.3.9.8 [ETM AMC A2 3.9.8] (1) Field acoustical calibrations All components of the system, excluding the windscreen, should be in place at this time, including cables, attenuators, gain and signal-conditioning amplifiers, filters (including pre-emphasis) and power supplies. During Page 128 ### Page 129 calibration, attenuators and gain stages should be set to prevent overload and to maintain the calibration signal level on the reference level range within the limits specified in A36.3.6.6 of part 36. If any switchable filters that could affect the calibration signal are utilized during measurements, then calibrations should be performed both with and without these filters enabled. Components of the electrical system should not be added, removed or replaced without re-calibrating the entire system immediately before and after each change. 401(c)(14) Windscreen loss adjustment AMC §A36.3.9.10 [ETM AMC A2 3.9.10] (1) Determination of windscreen data adjustments The physical condition of a windscreen can significantly affect its performance, and manufacturer-provided windscreen data adjustments for insertion loss are valid only for new, or clean, dry windscreens. For these adjustments, a single set of values based upon windscreen insertion loss tests at grazing incidence may be used when the angles of incidence of sound emitted from an aircraft are within +30° of grazing incidence. For other cases, the windscreen insertion loss adjustments should be determined and applied on the basis of intervals between angles tested not exceeding 30°. When the windscreen data adjustments provided by the manufacturer are presented in the form of curves, care should be taken to include the insertion loss throughout each one-third octave band, rather than just at the nominal midband frequency. Windscreen insertion loss can vary substantially within the frequency range of a single band and must be averaged or faired to more accurately correct one-third octa ve band data for the presence of the windscreen. Windscreen data adjustments may also be obtained by free-field calibration in an anechoic chamber. 401(c)(15) Measurement sy stem background noise AMC §A36.3.10.1 [ETM AMC A2 3.10.1] (1) Measurement system noise Since measurement system noise can add energy to m easured aircraft noise levels, the background noise measurement described in A36.3.10.1 of part 36 should be made with all gain stages and attenuators set as they would be used during the aircraft noise certification measurements. If it is expected that multiple settings will be required during the measurements, background noise data should be collected at each of these settings. Care should be taken to ensure that the background noise is truly representative of that present during the aircraft noise certification tests. (2) Mean background noise assessments At least 30 seconds of background noise data must be time-averaged to determine the mean level for each onethird octave band. The PNL value for this averaged spectrum should then be calculated using the procedures defined in A36.4.1.3 a) of part 36. The aircraft noise level data should also be analyzed and PNL values calculated for each spectral record. The maximum aircraft PNL value should be at least 20 dB above the PNL of the averaged background noise spectrum for the data to be considered acceptable. 401(d) Calculation of EPNL from Measured Data §A36.4 [ETM 4.1.4] 401(d)(1) Instantaneous sound pressure levels GM §A36.4.1.2 [ETM GM A2 4.2] (1) Instantaneous sound pressure levels For the purposes of this procedure, “instantaneous” sound pressure levels are considered to be one-third octave band sound pressure levels for each one-half second record obtained using a continuous exponential averaging Page 129 ### Page 130 process, as described in A36.3.7.5 of part 36 or its equivalent. 401(d)(2) Tone correction calculation AMC §A36.4.3.1 [ETM AMC A2 4.3.1] (1) Data precision for tone correction computation Prior to Step 1, it is recommended that all one-third octave band sound pressure levels be temporarily rounded to 0.1 dB resolution. The tone correcti on procedure presented here includes several steps that utilize decibel level criteria to characterize the significance of tonal content. These criteria can become artificially sensitive to small variations in level if a resolution finer than 0.1 dB is used in the computations. 401(d)(3) Adjustments relati ng to background noise AMC §A36.4.3.1(d) & (e) [ETM AMC A2 4.3.1 (Steps 4, 5)] (1) Data adjustments for background noise When the technical procedure presente d in 306(c)(2) of this AC is used for adjustment for the effects of background noise, Steps 4 and 5 of this tone correction procedure should be modified as follows: — Step 4. The LGB should be used in place of the highest frequency band (i = 24); and — Step 5. A new slope, s’(25,k), should be calculated for the band beyond LGB as described for an imaginary 25th band. This slope should be used in place of the slope derived from the actual level of the band beyond LGB. 401(d)(4) Data resolution afte r tone correction calculation AMC §A36.4.3.1(j) [ETM AMC A2 4.3.1 (Step 10)] (1) Data precision (after calculation of tone correction factor) At this point, the original sound pressure level resolu tion of 0.01 dB should be restored. Although the required precision of reported EPNL is 0.1 dB, all other intermedia te calculations external to the tone correction process should maintain a precision of at least 0.01 dB. (2) Identification of pseudo-tones Section 403(b)(2) presents guidance material on methods for identifying pseudo-tones. Note that the use of ground plane or 33 ft (10 m) microphones is supplemental to the required 4 ft (1.2 m) microphones and is allowed only for identification of frequency bands within which pseudo-tones might occur and not for the determination of aircraft certification noise levels. (3) Tone correction factor adjustment When tone correction f actors result from false or fictitious tones, recalculation is allowed using revised sound pressure level values, based on narrow-band analysis, of the smoothed spectral levels obtained in Step 7. Once the levels have been revised, the tone correction factor should be recomputed for the revised one-third octave band spectrum. This recomputed maximum tone correction factor should be applied, even if it occurs at or near the band associated with an artific ial tone, and approval by the FAA should be obtained for the methodology used. 401(d)(5) Band sharing adjustment GM §A36.4.4.2 [ETM GM A2 4.4.2] Page 130 ### Page 131 (1) Band-sharing adjustment concept The one-third octave band filtering process specified for analysis of aircraft noise certification data in A36.4.3.2 of part 36 may allow the tone correction procedure to und er-predict a tone correction factor when the frequency of a tone is located at or near the edge of one or more one-third octave bands. To account for this phenomenon, a band-sharing adjustment is computed that takes advantage of the fact that, as a result of the Doppler effect, a tone that is suppressed at PNLTM will probably appear normally in the spectra that occur before or after PNLTM. By averaging the tone correction factors calculated for the spectra within a 2-second period around PNLTM, the tone correction factor that would have occurred at PNLTM if it were not suppressed can be reasonably estimated. 401(d)(6) Calculation of band sharing adjustment AMC §A36.4.4.2 [ETM AMC A2 4.4.2] (1) Computation of band-sharing adjustment Although part 36 refers to identification of the frequency bands in which maximum tone corrections occur for the records near PNLTM, the presence or absence of band-sharing cannot be established merely by observing these frequencies. Even though the maximum tone that occurs in a one-third octave band spectrum may not be related to the band of maximum tone correction in the PNLTM spectrum, a related tone may still be present. Therefore, the average of the tone corrections of all spectra within one second (i.e., five, one-half-second data records) of PNLTM should be used regardless of the bands in which maximum tones are found. If the band-sharing adjustment is believed to result from effects other than band-sharing, the applicant should demonstrate its absence for each event. (2) Adjustment of PNLTM for band-sharing The band-sharing adjustment should be computed before the determination of the 10 dB-down period and should be included in the reported PNLTM and EPNL values for the test condition data. (3) Application of band-sharing adjustment for simplified procedure When the simplified procedure is used to adjust data to reference conditions, the band-sharing adjustment should be applied to the PNLTr at time of PNLTM before “∆1” and EPNLr are calculated. (4) Application of band-sharing adjustment for integrated procedure When the integrated procedure is used to adjust data to reference conditions, a new band-sharing adjustment should be calculated as in A36.4.4.2 of part 36. This new band-sharing adjustment uses the average of the tone correction factors of the PNLTMr spectrum and the tw o preceding and two succeeding spectra, after adjusting them to reference conditions, and should be applied to the PNLTMr value prior to identification of the reference condition 10 dB-down points and calculation of EPNLr. 401(d)(7) Noise Duration AMC §A36.4.5 [ETM AMC A2 4.5] (1) Noise duration (10 dB-down period) This period is the portion of the aircraft flyover in wh ich the measured noise level is within 10 dB of PNLTM (i.e., the period to be used for the calculation of EPNL). To ensure an adequate duration of recorded noise, recording systems should be activated, and the aircraft maintaining a stable condition, when the noise level at the first microphone location is estimated to be approximately 20 dB(A) below what is expected to be L Amax. Care should be taken during use of the flight path intercept method (see 402(a)(1)(i)) to ensure that noise levels have fallen 20 dB(A) below LAmax before flight path go-around procedures are initiated. ### Page 132 Note.— If recorded data do not encompass the entire 10 dB-down period, an EPNL cannot be calculated from those data, and the event should not be used for aircraft noise certification purposes. (2) Identification of the first and last records within the noise duration When identifying the records that define the limits of the noise duration, those records having PNLT values closest to the actual value of PNLTM – 10 dB should be used. As a result, the PNLT values for the PNLTM – 10 dB points may not always be greater than or equal to PNLTM – 10 dB. In order to illustrate the correct identification of the 10 dB-down points, Figure 9 provides examples of PNLT time-histories made up of records calculated from measured one-half-second values of SPL in accordance with the procedures specified in A36.4.2 of part 36. The shaded record k M represents the reco rd associated with PNLTM. Shaded records kF and kL represent respectively the first and last 10 dB-down points. In the first example the PNLT value associated with kF is greater than PNLTM – 10. The PNLT value associated with kL is less than PNLTM – 10. In the second example there are two records after kM with a value equal to PNLTM – 10. In this case kL is the last of the two records. The first 10 dB-down point, k F, is the record closest in valu e to PNLTM – 10, ignoring any records that precede it with greater values but which are less in value than PNLTM – 10. Note.— In all cases in the calculation of EPNL, the contribution of all the records from kF to kL inclusive should be included. 401(d)(8) Equations for co mputing duration correction GM §A36.4.5.4 [ETM GM A2 4.6.2] (1) Duration correction factor The equation for the duration correction factor, D, in A36.4.5.4 of part 36 is valid only for records of one-half second in length. The constant value 13 is used to normalize the one-half second values to the 10 seconds standard duration (i.e., 10 seconds duration comprising twenty 0.5-second data records and 10 log (20) = 13.01). ### Page 133 PNLTM –10 dB Tone corrected perceived noise level (PNdB) PNLTM Time kLkMkF Time kLkMkF PNLTM –10 dB PNLTM Tone corrected perceived noise level (PNdB) Figure 9. Illustrated example of identification of first and last 10 dB-down records 401(e) Reporting of Data to the FAA §A36.5 [ETM 4.1.5] 401(e)(1) Compliance records GM §A36.5.2 [ETM GM A2 5.1] (1) Compliance records For compliance with A36.5 of part 36, all data measured during noise certification testing, including time-histories of physical measurements, noise recordings, instrument calibrations, etc., are to be recorded in permanent form and made available to the FAA for review, inspection and approval. A common procedure is for the applicant to submit representative samples of test data for each noise measurement point and adjustments to measured data to permit the FAA to determine compliance with part 36. The applicant may either submit the complete test records along with the required data adjustments, or when approved by the FAA, the applicant may instead submit samples of test data along with the required data adjustments. 401(e)(2) EPNL average values GM §A36.5.4.2 [ETM GM A2 5.4] (1) Average EPNLr levels when using the NPD equivalent procedure Page 133 ### Page 134 For airplanes the average value of EPNL r from an NPD database (see 402( a)(1)(ii)(A)) is the noise level determined along the regression line through the adjusted data set at the appropriate thrust (power) and distance values, including any other additional adjustments necessary (e.g., adjustment to the aircraft reference speed). (2) Single test values When more than one noise measurement system is used at any one noise measurement point, the resulting noise level is to be the average of the measured noise levels for each noise measurement po int. This requirement does not apply to noise levels measured by microphones not required for acquisition of noise certification data. (3) Valid conditions All valid noise measurements are to be included in the confidence interval calculations even when they produce results that are outside the 90 percent confidence limit of ±1.5 dB. The cause of erratic or possibly invalid noise data may include testing under different temperature and humidity extremes, anomalous winds aloft, changes in noise measurement system components, changes in aircraft hardware, background noise, shift in instrument calibrations, or not testing in accordance with the approved test plan, etc. The FAA is to make a determination, during the course of noise certification testing, as to the validity of all noise measurements. A noise measurement may not be excluded from the confidence interval calculations at a later date without approval by the FAA. Noise measurements determined in the field to be invalid for any reason may need to be repeated in order to achieve the required minimum number of valid test runs. 401(e)(3) Calculation of 90 percent confidence intervals AMC §A36.5.4.2 [ETM AMC A2 5.4] (1) Methods for calculating 90 percent confidence intervals Section 305, provides confidence interval calculation methods for clustered measurements, regression mean line, static-test-derived NPD curves and analytically derived NPD curves, along with wo rked examples. Calculation methods for determining 90 percent confidence interval values for clustered and pooled data sets are presented in 305(f)(2) and 305(f)(5) of this AC. (2) Retest requirements The FAA may require an applicant to retest or provide additional test data for any of the three noise measurement points when the reported results indicate: a) a required measurement is reported to be invalid; b) an insufficient number of measurements were conducted by the applicant to determine a suitable data sample; c) data scatter indicates that the data are not from a normal population or trend (e.g., a discontinuity due to low power SBV operation); d) the 90 percent confidence interv al for a noise measuring condition exceeds the allowable ±1.5 dB; or e) the test was not conducted in accordance with an approved noise certification compliance demonstration plan. 401(f) Nomenclature: Symbols and Units §A36.6 [ETM 4.1.6] (Reserved) Page 134 ### Page 135 401(g) Sound Attenuation in Air §A36.7 [ETM 4.1.7] (Reserved) 401(h) Adjustment of Helicopter Flight Test Results §H36.205 [ETM 4.1.8] The objective of a noise certification test is to acquire da ta for establishing an accurate and reliable definition of a helicopter’s noise characteristics. Appendix H establishes a range of test conditions and procedures for adjusting measured data to reference conditions. 401(h)(1) Adjustments to reference conditions GM §H36.205(a) [ETM GM No. 1 A2 8.3] (1) Adjustments to reference conditions Most noise certification tests are conducted during cond itions other than the reference conditions. This includes differences in height, lateral positio n, airspeed, rotor speed, temperature and relative humidity. Therefore, measured noise data should be adjusted to reference conditions to determine whether compliance with the noise certification limits of Appendix H may be achieved. Both positive and negative adjustments must be applied for the differences between the test and reference conditions. Adjustment procedures and analysis methods should be reviewed and approved by the FAA. The FAA should ensure that data adjustment and analysis methods that are proposed by applicants satisfy the requirements of part 36 and approved procedures. Any changes, including software revisions, firmware upgrades or instrumentation changes, are subject to the review of the FAA before they can be used for noise certification evaluations. Program validation should be planned and the required information submitted to the FAA early in the certification cycle, since th e time required for evaluation and approval may vary depending upon the issues encountered. (2) Non-positive SPLs Whenever non-positive one-third octave band aircraft noise levels are obtained, whether as part of the original one-third octave band analysis or as a result of adjustments for background noise, or other approved procedures, their values should be included in all relevant calculations. The practice of “band-dropping”, where masked levels are methodically set equal to zero, is not considered to be an acceptable substitute for reconstruction of masked levels as per the background noise adjustment guidance provided in 306. For any aircraft noise spectrum subject to adjustment to reference conditions, all one-third octave bands, including those containing masked levels or reconstructed levels, including values less than or equal to zero dB, should be adjusted for differences between test and reference conditions. (3) Direction of flight considerations Since flyovers are made in two directions with headwind and tailwind components, the lateral (sideline) microphones will be either “left sideline” or “right sideline” depending on the direction of flight. Hence sideline flyover data need to be sorted by left microphone and right microphone for data adjustments and reporting. Note that sorting by left and right sideline microphone is also appropriate for takeoff and approach if more than one direction of flight is used. It should also be noted that an equal number of flyover test runs with headwind and tailwind components are required. If after analysis the applicant finds that there is at least the required minimum of three measured values in each flight direction, but there are more in one direction than in the other, the applicant then will need approval by the FAA as to which are to be used in the determination of the final EPNL value for flyover. ### Page 136 401(h)(2) Reference data sources GM §H36.111(c) [ETM GM No. 2 A2 8.2.1] (1) Manufacturer’s data Adjustment of noise values from test to reference cond itions should be based on approved manufacturer’s data. Manufacturer’s data should include: a) reference flight profiles; b) takeoff and flyover engine power settings at reference conditions; and c) reference airspeeds. 401(h)(3) Adjustments to measured noise data GM §H36.205 [ETM GM A2 8.3.1.2] (1) Reference flight path noise propagation angle In calculating the position of the PNLTM on the reference flight path, the acoustic emission (i.e., noise propagation) angle () relative to the flight test path must be kept the same as for the test flight path. The elevation angle () relative to the ground plane is not constrained, and determination and reporting of this angle is required. (2) Maximum adjustments To prevent excessive adjustments to the measured data , the summation of all the adjustments for differences between the test flight path and the reference flight path for flyover and approach is limited to 2 EPNdB. For takeoff the summation of the adjustments is limited to 4 EPNdB of which the sum of 1 and the –7.5 log term from 2 must not exceed 2 EPNdB. The additional allowance for takeoff acknowledges th at larger differences between the test flight path and reference flight path can occur for this condition as a result of the influence of wind speed on the test flight path. It is recommended, however, that the applicant note that methods discussed in H36.3(c)(a)(c)(2) can be used to minimize this difference for takeoff. 401(h)(4) Takeoff profile GM §H36.205(b) [ETM GM A2 8.1.2.1] (1) Reference takeoff profile Figure 10 illustrates the reference takeoff profile and an idealized test or measured takeoff profile under zero wind conditions. ### Page 137 20 m (65 ft) 500 m (1640 ft) 10/12/2017 AC 36-4D L   +  A F C Sstarboard Cr Lr r Fr Ir Reference take-off path projection Reference take-off path Measured take-off path Sport 20 m (65 ft) 500 m (1 640 ft) 150 m (492 ft) 150 m (492 ft) Figure 10. Comparison of mea sured and reference takeoff profiles The reference takeoff profile is a straight line segment. It starts from a defined point C r that is 1640 ft (500 m) from the center microphone location A an d at a height of 65 ft (20 m) above the ground. The reference climb angle ( β) of the straight line path will depend on the certificated best rate of climb and V y at the reference conditions. The reference profile ends at a point I r which will encompass the 10 dB-down period of the noise measurements. Note.— For clarity the location of the test and reference PNLTM points, L and Lr, are illustrated at the same position for both the centerline noise measurement point A and the starboard lateral (sideline) noise measurement point S. Normally however L, and hence Lr, will be a different position on the test and reference flight paths for each noise measurement point. (2) Reference climb angle The reference climb angle, β, is based on the best rate of climb and V y airspeed determined from approved manufacturer’s data for the takeoff performance of the helicopter at the reference conditions. Since airspeed is defined as being in the direction of the flight path, the climb angle β is the arcsine of the ratio of best rate of climb to Vy. On a helicopter that is engine-power-limited at the reference conditions, the best rate of climb has to be calculated from the minimum specification engine(s) performance. On many helicopters the takeoff characteristics will be dependent on gearbox torque limit, and this will be typically less than the torque associated with minimum specification engine(s) at the reference conditions. Since all procedures have to be consistent with the airworthiness regulations, the gearbox takeoff torque limit should be used to calculate the applicable best rate of climb at the maximum noise certification weight fo r those helicopters which are performance-limited by the gearbox characteristics at the reference conditions. 401(h)(5) Takeoff test conditions AMC §H36.101(d) [ETM AMC No. 1 A2 8.1.2.1] (1) Takeoff requirements The takeoff profile is commenced from a level flight at a height of 65 ft (20 m). After reaching position C, takeoff power has to be applied to initiate the climb. The takeoff power will either be dependent on the gearbox torque Page 137 ### Page 138 limit for takeoff or minimum installed engine(s) takeoff power torque at the reference conditions at sea level and 77ºF (25C). (2) Test airspeed The best rate of climb airspeed V y to be used is that determined from the takeoff performance at sea level and 77F (25C) during airworthiness certification. This is to be maintained during the complete takeoff procedure. To account for test-to-test variation and slight variations during each test run, a tolerance of ±5 kt (±9 km/h) is allowed. (3) Rotor speed The mean value of the rotor speed during the 10 dB-down period is to be within ±1 percent of the maximum normal operating rotor speed value at the reference takeoff condition. (4) Flight path deviations To minimize lateral flight path deviations, and hence the difference in noise levels due to off-track position at the PNLTM emission point, the helicopter must fly over the reference flight track during the 10 dB-down period within ±10 or ±65 ft (±20 m) from the vertical, whichever is the greater. This is illustrated in Figure 11. There is no direct height limitation, but the adjustments that take into account differences between the reference and test sound propagation distances at PNLTM are limited to 2 EPNdB as discussed in H36.205(f) and H36.111(c) & (d) of part 36. Lateral cross-sectional view of take-off corridor 20 m 20 m Reference flight track 10° 10° Ground surface Figure 11. Lateral deviation tolerances for takeoff (5) Helicopter test weight Page 138 ### Page 139 The weight of the helicopter during the noise certificatio n demonstration (see H36.101(b) of part 36) must lie within the range of 90 percent to 105 percent of the maximum takeoff weight for the takeoff demonstration. No adjustment of the noise data to maximum takeoff weight is required. At least one takeoff test run must be completed at or above this maximum certificated takeoff weight. If the value of the maximum takeoff weight selected for noise certification is less than that used fo r airworthiness certification, then the lower weight may become the operating limitation defined in the appropriate section of the RFM 401(h)(6) Takeoff flig ht test procedures AMC §H36.103 [ETM AMC No. 2 A2 8.1.2.1] (1) Test takeoff profile The test takeoff profile requires stabilized flight conditions only over the 10 dB-down period in the climb portion of the procedure. (2) Number of test runs At least six test runs are required with simultaneous noise measurements at each of the noise measurement points. It should also be remembered that synchronized noise and flight path data are required. Since it cannot be determined until the analysis is partly completed if each test run meets all the requirements of Appendix H, the applicant will find considerable merit in conducting additional takeoff test runs. Experience suggests that 8 to 10 test runs would normally provide an adequate safeguard against some test runs being determined invalid during subsequent analysis. If additional test runs are conduc ted and more than six valid noise measurements are simultaneously obtained at all three measurement points, then the results of such test runs are also required to be included in the averaging process fo r calculating EPNL. The results of test runs without simultaneous noise measurements at all three measurement points are not included in the calculation process. (3) Flight airspeed tolerance A ±5 kt (±9 km/h) tolerance about the reference airspeed is specified in H36.103(b) of part 36. This is not intended to allow tests at different speeds but rather to account for variations during the 10 dB-down period which occur during an individual test run as a result of the pilot attempting to maintain the other takeoff requirements and testto-test variations. The value of Vy is published in the takeoff performance section of the RFM and is typically defined as an IAS. The applicant should note the reference airspeed is the tr ue airspeed (TAS). Since most airspeed instruments do not indicate the TAS value, airspeed calibration curves and meteorological conditions should be used to convert between TAS and IAS. (4) Horizontal adjustment of climb initiation Position C in Figure 10 may be varied, subject to approval by the FAA, to minimize the difference between the test and reference heights vertically above the flight track noise measurement point. This difference can result from the effect of wind on the climb angle during testing. Figure 12 illustrates the case of a headwind. Note that even for zero or very low wind, the transition from the horizontal flight to the climb can take a significant time. This will be the case normally on larger and heavier helicopters. The resulting flight path could be well below the reference profile. In this case there would be merit in moving position C further away from the noise measurement point. This is illustrated in Figure 13. (5) Vertical adjustment of climb initiation Subject to approval by the FAA, the height of the initial level flight may also be varied in order that the height (distance) adjustment associated with the climb phase can be minimized. This is an equivalent procedure, which Page 139 ### Page 140 can be used in the place of adjusting the horizontal location of position C fr om the flight track noise measuring point to achieve the same result. The applicant should note that under many test conditions no such adjustments are required to comply with the data adjustment procedures defined in H36.205(f). Fr Cr C F Take-off test path with adjustment to position “C” Take-off test reference profile Position “C ” is adjusted to p r osition “C” to reduce the height difference relative to the reference height over the measurement point Position “C” is moved towards measurement point “A” Wind direction A Figure 12. Adjustment of takeoff profile position “C” for headwind Note.— The above procedures for horizontal or vertical adjustment of climb initiation are based on consideration of the height above the flight track or center noise measuring point, even though the adjustments to the noise measurements are applied for the PNLTM point. However, since the PNLTM point, which is normally within close proximity of the overhead point, cannot be determined until after the noise analysis is conducted, use of height over the noise measuring point to determine the location of position C (or the initial horizontal height) is acceptable. (6) Practice flight Irrespective of which method is used to control the height over the flight track noise measurement point, an applicant may find it helpful, if not essential, to conduct a number of practice or pre-noise certification test runs to adjust the height/location of position C. With prior ap proval of the FAA, these practice runs can be excluded from the noise compliance evaluation. These runs should also be documented in the noise certification report as practice flights. (7) Power setting Takeoff power at sea level and 77°F (25 C) has to be applied at position C to initialize the climb. On many helicopters the airworthiness power limit will be set by the takeoff gearbox torque limit. When this is not the case, the takeoff torque will be that torque determined during the airworthiness certification and will be based on the minimum specification engine(s) power. ### Page 141 Cr Fr Ir F A C CA > C Ar At zero or low wind conditions Reference take-off/climb path Resulting take-off path with adjustment at position “C” At zero or low wind conditions, position “C” is adjusted to reduce height difference relative to reference height over noise measurement station “A” Position “C” is moved further away from noise measurement station “A” Figure 13. Adjustment of takeoff prof ile position “C” for zero or low wind In some cases the applicant may find that the takeoff gearbox torque limit to which the helicopter is to be airworthiness certificated has not been approved and hence cannot be used during the noise certification test. When testing only at a lower torque is possible, the FAA may approve, as an equivalent procedure, the extrapolation of noise data from lower torque settings. Tests conducted at the maximum available, and a minimum of two, lower gearbox torque settings would be required for extrapolation, subject to approval by the FAA, to a higher torque value. Experience suggests that extrapolation of no more than 10 percent is likely to be acceptable. The applicant will also need to document in detail the extrapolation procedure to be used. (8) Rotor speed Rotor speed may be manually or automatically varied on some helicopters. On many designs variation in rotor speed can occur due to the limits of the engine/rotor governing system. In order that the noise levels are representative of normal takeoff operation, the rotor speed is to be the maximum normal value associated with the reference takeoff airspeed. Since on most helicopter s small rotor speed changes occur during a stabilized flight, a ±1 percent rpm variation in the rotor speed is allowed. Note.— Noise measurements should be made at the maximum rotor speed during normal operations. Testing at the maximum tolerance rpm is not required. On some helicopter designs more than one rotor speed ma y be available (see 402(e)(1)(vi)). If multiple rotor speeds can be used for normal operations, then noise certification has to be conducted at the highest value allowed at the reference conditions. If the highest speed is limited to special operations, or if the helicopter is configured such that the highest rotor speed cannot be used at the reference conditions or test height, then, subject to approval by the FAA, testing at a lower rotor speed may be allowed. On some helicopter designs, rotor speed may be automatically varied within the 10 dB-down period. In such cases, a reference rotor rpm schedule as a function of position along the reference flight path should be defined and tests should be conducted so as to maintain the test rotor rpm within 1% of the reference rpm schedule. If the variation in rotor speed results in changes to the best rate of climb, a non-linear reference flight profile should be defined and used in the calculation of reference dist ances for adjustments of the noise data to reference conditions. For example, for a helicopter that automatically varies rotor speed (N r) during takeoff, a reference rotor speed Page 141 ### Page 142 schedule may be defined as a function of height above ground level along the reference flight path during the 10 dB-down period as illustrated in Figure 14. In the typical case where the helicopter is main gearbox torque limited at the noise certification reference conditions, the rotor speed schedule may also result in a non-constant (curved) reference flight path segment during the N r transition as illustrated in Figure 14. The 1% test requirement for rotor speed would be applied to the reference rotor speed schedule as shown in Figure 14. 86 88 90 92 94 96 98 100 102 104 –600 –500 –400 –300 –200 –100 0 100 200 300 400 500 600 Horizontal distance from flight track measuring station (m) Rotor RPM (% N )r Height AGL (m) 0 40 80 120 160 200 240 280 320 360 400 Rotor rpm (% N )r Height AGL (m) Constant reference flight path segment Non-constant reference flight path segment Constant reference flight path segment Approximate 10-dB down period Reference N schedule r End of N transition r Beginning of N t r a n s i t i o nr –1% Nr –1% Nr +1% Nr +1% Nr Figure 14. Example of a reference takeof f flight path and rotor speed schedule (with ±1% Nr limits) for a variable rotor speed helicopter (9) Flight path guidance To meet the requirement of being within ±10 of the vertical, the applicant may need to use flight track markings that are clearly visible and/or on-board flight track guidance instrumentation and a real-time position measuring system or some other approved method of checking that this requirement is satisfied. Pilot visibility of the ground when climbing may be somewhat limited and thus markers well ahead of the noise measuring point may be required. Some of the flight path measurement systems outlined in 302, provide flight track data in real time, or in a short period after the test run. In this case the applicant will readily be able to determine if a test run is within the allowable deviation limits. If a simpler system is utilized, such as photographic scaling based on the use of still cameras, the applicant may find it useful, if not essential, to develop a method to enable timely confirmation that the test run is acceptable. The actual height and off-track deviations do not need to be established at the time of the test run. However the applicant needs to ensure that otherwise acceptable test runs are not rejected during analysis for failing to meet the ±10 limit for lateral deviation. 401(h)(7) Flyover configuration GM §H36.105 [ETM GM No. 1 A2 8.1.2.2] (1) Reference flyover profile Appendix H specifies the reference procedure as a level flyover at 492 ft (150 m) above the ground at the flight track measurement point as illustrated in Figure 15, in which the reference flight profile is indicated as D r to Jr and the test profile as D to J. ### Page 143 G S(sideline) A S Reference overflight path projection Measured overflight path 150 m (492 ft) 150 m (492 ft) 150 m (492 ft) Dr Mr Gr Jr JD r  M Reference overflight path Figure 15. Comparison of mea sured and reference flyover profiles The reference airspeed is 0.9 V H, 0.9 V NE, 0.45 V H + 65 kt (0.45 V H + 120 km/h) or 0.45 V NE + 65 kt (0.45 VNE + 120 km/h), whichever is less, throughout the 10 dB-down period. The rotor speed (rpm) is fixed at the maximum normal operating value. Note that if VH is greater than VNE then the reference airspeed will be related to VNE. Note.— For clarity the location of the test and reference PNLTM points, M and Mr, are illustrated at the same position for both the centerline noise measurement point A and the lateral noise measurement point location S. Normally, however, M, and hence Mr, will be a different position on the test and reference flight path for each noise measurement point. GM §H36.105 [ETM GM No. 2 A2 8.1.2.2] (1) Flight path deviations To enable the flyover noise characteristics to be obtained, the flyover test has to be a level flight at a fixed height above the flight track noise measurement point. The test runs also have to be within ±10  or ±65 ft (±20 m), whichever is the greater, from the vertical throughout the 10 dB-down period. The ±65 ft (±20 m) is not relevant in the case of flyover since the off-track deviation allowed, at the test height, is controlled by the ±10 requirement. (2) Test airspeed tolerance The flight airspeed is defined in H36.3(d) of part 36 and a ±5 kt (±9 km/h) tolerance from the reference airspeed during each flyover is allowed within the 10 dB-down period. The power is to be stabilized, and the mean value of the rotor speed during the 10 dB-down period is to be within ±1 percent of the normal operating rpm value for each flyover. (3) Source noise adjustment testing The applicant should consider the requirement for source noise adjustments since it is unlikely that the tests can Page 143 ### Page 144 be conducted precisely at the reference temperature of 77F (25C), reference rotor speed and reference airspeed. This dictates that, if the test advanc ing blade tip Mach number is different from the reference Mach value, the development of a PNLTM versus advancing blade tip Mach number sensitivity curve is necessary. This requires testing at different flight speeds around the reference flight speed. The number of additional test runs will to some extent depend on the character of the variation of PNLTM with flight speed, but since this cannot be determined until after the analysis is complete, conservative estimates for the number of additional test runs and the actual airspeeds to be used need to be considered. Note.— The equivalent Mach number procedure, discussed in section 10 of 401(h)(8), is an acceptable method of compliance that eliminates the need for a source noise adjustment. (4) Helicopter test weight The weight of the helicopter during the noise certification demonstration (see H36.101(b)(6)(ii) of part 36) must lie within the range of 90 percent to 105 percent of the maximum takeoff weight for the flyover demonstration. At least one flyover test must be completed at or above this maximum certificated weight. If the value of the maximum takeoff weight selected for noise certification is less than that used for airworthiness certification, then the lower weight may become the operating limitation defined in the appropriate section of the RFM. 401(h)(8) Flyover test procedure AMC §H36.105 [ETM AMC A2 8.1.2.2] (1) VH VH is defined as the airspeed in level flight at the reference conditions and maximum certificated takeoff weight and is obtained using the minimum specification engine(s) torque at maximum continuous power. V H will need to be determined specifically for the noise certification fl yover tests, since its determination is not required for the airworthiness certification. V H, by itself, is never limited by airworthiness considerations. However the maximum continuous power on which it is based may be limited due to airworthiness issues and that, in effect, could limit the value of VH. (2) VNE VNE is determined as a part of the airworthiness approval and will therefore be readily available. (3) Reference airspeed On some helicopters VH may be in excess of the level flight V NE imposed and approved by the FAA. The intent of noise certification is not to relate test airspeeds to reference airspeeds that may be beyond the airworthiness VNE limit of the helicopter. Under H36.3(d) of part 36, V NE would therefore apply in place of VH. Also on some helicopters with high airspeed capabilities, the typi cal cruise airspeed will be less than 0.9 V H (or 0.9 VNE) and thus if 0.9 VH (or 0.9 VNE) were used as the reference, it would no longer be representative of a cruise flight. In this case a lower airspeed of 0.45 VH + 65 kt (0.45 VH + 120 km/h) or 0.45 VNE + 65 kt (0.45 VNE + 120 km/h) is used. This applies when 0.9 VH (or 0.9 VNE) is 130 kt (240.8 km/h) or higher (i.e., when VH (or VNE) is 144.4 kt (267.5 km/h) or higher). Thus the reference airspeed will be the least of the following four airspeeds: a) 0.9 V H; b) 0.45 V H + 65 kt (0.45 VH + 120 km/h); c) 0.9 V NE; or d) 0.45 V NE + 65 kt (0.45 VNE + 120 km/h). The text continues at https://openairdocs.org/ac/36-4.md?from=145 --- From OpenAirDocs.org, an independent open copy of FAA data. Only the Federal Register edition on govinfo.gov is the official legal text of an AD.