Aircraft noise reduction technology

(NASA-TM-X-68241) AIRCRAFT NOISE N74-10917 REDUCTION TECHNOLOGY (NASA)

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G3/02 _ 2 1648

NASA TENC ICAL-

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AIRCRAFT NOISE

REDUCTION TECHNOLOGY

Lewis

Research Center

Cleveland, Ohio

March 1973

REPRODUCED BY NATIONAL TECHNICAL INFORMATION SERVICE U.S. DEPARTMENT OFCOMMERCE SPRINGFIELD, VA. 22161 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.

NASA TM X-68241 5. Report Date 4. Title and Subtitle March 1973 AIRCRAFT NOISE REDUCTION TECHNOLOGY 6. Performing Organization Code 8. Performing Organization Report No.

7. Author(s) E-7397 10. Work Unit No.

9. Performing Organization Name and Address Lewis Research Center Lewis Research Center 11. Contract or Grant No.

National Aeronautics and Space Administration Cleveland, Ohio 44135 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code 14. Sponsoring Agency Code Washington, D. C. 20546 15. Supplementary Notes 16. Abstract This report, describing the NASA noise reduction technology programs and plans, was prepared for use by the Environmental Protection Agency in the aircraft/airport noise study.

Separate sections of the report deal, respectively, with characterizing the effects of aircraft noise on individuals and communities; the status of aircraft source noise technology; opera- tional procedures to reduce the impact of aircraft noise; and aspects of NASA relations with the military services in the aircraft noise area. The report is a summary and guide to aircraft noise research and technology and includes references to more detailed technical literature on the subjects discussed.

17. Key Words (Suggested'by Author(s)) 18. Distribution Statement Aircraft noise; Sonic boom; Noise abatement Unclassified - unlimited procedures; Aerodynamic noise; Powered-lift aircraft noise; Noise impact characterization; Propulsion-system noise 19. Security Classif. (of this report) 20. Security Classif. (of this page) Unclassified Unclassified * For sale by the National Technical Information Service, Springfield, Virginia 22151

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Preceding

FOREWORD The Environmental Protection Agency has requested the participation of the National Aeronautics and Space Administration in a comprehensive study of the noise problems resulting from aircraft operations.

This study is being conducted by EPA in consultation with federal, state, and local agencies and other interested persons.

The results of the study will be contained in an EPA report to be submitted to the Congress as directed by the Noise Control Act of 1972 (PL 92-574).

As a part of the support requested, NASA has prepared this preliminary report for use by EPA in the aircraft/airport noise study. The intent of this report is to describe the NASA noise reduction technology programs and plans.

To put the NASA program in context, the general status of noise reduction technology is described.

The major problems to be overcome in reducing the noise generated by the various noise sources are discussed. The gains that have been made in noise technology are included, and the associated penalties in performance and the economic penalties are provided where known.

It must be emphasized, however, that this is chiefly a technology report and that the practicability of using this technology in actual aircraft systems involves detailed consideration of performance and economics for each individual application.

The EPA Task Group topics in this report are impact characterization, source abatement technology, flight operations, and military aspects.

Each topic is addressed in a separate section.

Section I of the report discusses the problem of characterizing the impact of air- craft noise on the individuals and communities affected.

The capability of various parameters and methods for accurately estimating the psychoacoustical impact on individuals and communities is assessed.

The difficulties associated with the effective use of the various impact characterizations are discussed.

The unique characteristics of the sonic boom and the interpretation of its impact on the community is reviewed.

In section II the broad subject of noise abatement technology is covered. The status of noise technology in the various propulsion system component areas is described. The relative importance of the component noise sources for the different aircraft applications, subsonic, supersonic, and powered lift aircraft, is discussed.

The results of the NASA Quiet Engine Program are presented.

The direction of future NASA source noise reduction research is also indicated.

Section III is concerned with the use of operational procedures to reduce the impact of aircraft noise. Primary emphasis is on the current NASA Two-Segment Approach Program. The potential noise benefits are indicated, and the results to date are described.

Future plans for this activity are also given.

iii Finally, section IV deals with various aspects of the NASA relation with the military services in the aircraft noise area. Formal and informal agreements, mutual programs, and exchange of personnel are all a part of the effort to realize the maximum noise re- duction benefits in the civil area from military research and development programs.

The NASA noise technology program is guided by a continuing study of various air- craft and propulsion systems, including conventional subsonic aircraft, supersonic air- craft, and powered-lift aircraft for short-haul applications. These studies are con- ducted in house as well as through contracts with industry and are often based on as- sumptions of future technology levels. Although quite useful as a guide to research and technology, the precise results of such studies are speculative.

Therefore, specific details of the various studies that NASA has conducted do not appear here, although general trends are sometimes included.

The coauthors of and the contributors to this report are as follows: NASA Headquarters Lee D. Goolsby William H.

Roudebush James J. Kramer Raymond P. Whitten NASA Ames Research Center Dallas G. Denery Clark White David H. Hickey NASA Langley Research Center Jimmy M.

Cawthorne Lucio Maestrello Andrew B. Connor Domenic J. Maglieri Latham Copeland Arnold W.

Mueller Philip M. Edge, Jr.

Tony L. Parrott David A. Hilton Robert J.

Pegg Robert H. Hosier Goldie C. Smith Harvey H. Hubbard NASA Lewis Research Center Carl C. Ciepluch Gene L. Minner E.

William Conrad Francis J. Montegani Robert J. Denington Leonard J. Obery Robert G. Dorsch William A. Olsen Charles E.

Feiler Edward J.

Rice John F. Groeneweg Nick E. Samanich William L. Jones James R. Stone Eugene A. Krejsa Uwe H. von Glahn James P. Lewis iv

CONTENTS

Page FOEOD...............................................111I FOREWORD I. IMPACT CHARACTERIZATION ANALYSIS ....................

I AIRCRAFT NOISE MEASUREMENTS ......................

I SUBJECTIVE RESPONSE ...........................

I MEASUREMENTS OF NOISINESS ...

........................

COMMUNITY REACTION TO AIRCRAFT OPERATIONS ................

SONIC BOOM ..................................

RESEARCH PLANS .........

...................

LABORATORY STUDIES ..... ...........................

AIRPORT COMMUNITY NOISE STUDIES ...................

SONIC BOOM STUDIES.. ............................

REFERENCES.....

........

...................

II. SOURCE ABATEMENT TECHNOLOGY ..... .................

SUBSONIC TRANSPORT AIRCRAFT ......................

PROPULSION SYSTEM NOISE .............

...........

Fan Noise ..........................................

Jet Noise ..... .....

........... ................

Other Noise Sources ... ..............................

Suppression of Internal Engine Noise ...................

Engine System Noise ............................

NASA Quiet Engine Program.

.......................

NASA Refan Program ............................

AIRCRAFT NONPROPULSIVE (AIRFRAME) NOISE ................

SUPERSONIC TRANSPORT AIRCRAFT .....

.....................

JET NOISE ABATEMENT TECHNOLOGY ..................

Variable Engine Cycles...

............

Suppression Devices ..

.. .....................

..

In-Flight Effects on Suppression Devices . . . . . . .

. . . . . . . . . .

Core Noise .............................

SONIC BOOM .................................

..........

Sonic-Boom Exposures For Steady Flight and a Steady Atmosphere . . .

Soi Bo Mnmiaio..................................

Sonic Boom Minimization . .

. . . . . . . . . . . . . . . . .

. . . .

. . 76 Effects of the Atmosphere .........................

Effects of Airplane Maneuvers . . . . .

. . . . . . . . . .........

Low Supersonic Operations . . . . . . . .. . . . . . . . . . . . .

. . . .

. 81 Summary of Results ............

.....................

V Planned Research..........

............................

POWERED LIFT AIRCRAFT . . . . . . . . . . . . . . . .

.

. . . . . . .

AUGMENTOR WING NOISE.......................

. 89 EXTERNALLY BLOWN FLAP NOISE ....

................

QCSEE PROGRAM ................................

NASA SHORT-HAUL AIRCRAFT SYSTEM STUDIES .................

ROTORCRAFT .................................

ENGINE NOISE *.

. .

..............

...................

ROTOR NOISE .................................

PLANNED RESEARCH.

...

..... ....................

GENERAL AVIATION AIRCRAFT .............................

QUIET VEHICLES.......................

........

QUIET PROPELLERS .. *.*..............................

REFERENCES .................................

BIBLIOGRAPHY ................................

III. OPERATING PROCEDURES FOR AIRCRAFT NOISE REDUCTION ........

PROGRAM HISTORY .......................

.............

TWO-SEGMENT APPROACH.

. *............................

STEP A: DEVELOPMENT AND FLIGHT EVALUATION OF A SPECIAL PURPOSE GLIDE SLOPE COMPUTER IN A BOEING 727-200 AIRCRAFT.

. * ...........................

STEP B: DEVELOPMENT AND FLIGHT EVALUATION OF TWO- SEGMENT AVIONICS USING THREE-DIMENSIONAL AREA NAVIGATION FOR GUIDANCE IN A DC-8-61 . ..............

STEP C: STUDY TO DETERMINE THE APPLICABILITY OF THE TWO- SEGMENT APPROACH TO ADDITIONAL JET TRANSPORTS ........

STEP D: STUDY TO DETERMINE THE SUITABILITY OF THREE- DIMENSIONAL AREA NAVIGATION TO PROVIDE VERTICAL GUIDANCE ...................................

STEP E: STUDY TO DETERMINE THE IMPACT OF THE TWO-SEGMENT APPROACH ON ATC . .....................

OTHER TECHNIQUES FOR NOISE ABATEMENT ...

............

FLIGHT TEST OF NOISE ABATEMENT APPROACHES USING A MICROWAVE LANDING SYSTEM ..........................

FLIGHT EVALUATION OF CURVED APPROACHES FOR NOISE ABATEMENT .

..........

.........

.........

NOISE ABATEMENT USING DECELERATING APPROACHES ........

ANALYSIS . ..............

.........

.. ......

Piloted Simulation .............................

vi Flight Test . . .. .. .

. . . . . . . . . . . .. . ... . . . .. ... .

AERODYNAMIC NOISE ............................

REFERENCES .................................

IV. MILITARY ASPECTS ...............................

TURBINE ENGINE PROGRAMS ........................

STOL AIRCRAFT PROGRAMS ROTORCRAFT PROGRAMS ..........................

vii

I - IMPACT CHARACTERIZATION

ANALYSIS

Basic to the effective control of aircraft noise and sonic boom is the development and validation of units of measurement that properly describe human response to them. Air- craft noise is of concern during ground operation, takeoff, climbout, and landing; whereas, sonic boom is of concern during the supersonic portion of flight over inhabited areas.

Included herein are brief discussions of NASA contributions to both noise and sonic boom characterization technology, the current state of the art, and some plans for future research. Specific mention is made of the research approaches used, the type of results obtained, and the manner of application of such information in aircraft noise control and its broader significance for community noise considerations.

AIRCRAFT NOISE MEASUREMENTS SUBJECTIVE RESPONSE Accurate measurement units describing human response are needed to define noise specifications for quieter aircraft and for aircraft noise certification. Units which are simple in concept and easy to use are also needed in predicting community annoyance and complaint patterns, land-use planning near airports, and airport traffic monitoring and control.

The definition of the measurement units is influenced by the psychophysiological characteristics of people, as seen by their responses to noise; the physical character- istics of the aircraft noise stimuli; and aircraft operational data, which include airport traffic, preferential runways, mix of aircraft types, and flight scheduling (time of day); such community environment considerations, as background noise levels; economic, geographic, and demographic factors; and community activities.

NASA has supported studies to characterize and evaluate individual and community response to aircraft noise. These efforts have been involved in developing an under- standing of how one perceives noise in degrees of noisiness and annoyance.

Individual response to single noise producing events and the overall responses of communities to actual commercial aircraft operations were studied.

This work has been accomplished largely under contract (refs. 1 to 8) by such organizations as Bolt, Beranek, and Newman; Tracor, Inc.; Boeing Vertol; and Stanford Research Institute, and has been coordinated with other government agencies and the National Research Council through the active participation of CHABA.

The effects of noise on people may include annoyance, speech interference, sleep interference, degradation of task performance, and hearing losses. Although each of these effects can be important in particular circumstances, annoyance and sleep inter- ference are judged to be of particular importance in understanding airport community situations and are thus focal points for NASA supported field and laboratory research in this area.

In a global sense it is recognized that all the above effects may contribute in some manner to shaping community attitudes; thus attempts have been made in a broad sense to correlate responses to noise exposure through the use of opinion poll surveys.

In the development and evaluation of measurement units, NASA studies used several test situations where environmental control was considered optimum. For example, many test situations involve a small anechoic room. A loudspeaker system is provided for playback of aircraft noise signals, and persons are arranged in a manner suitable for obtaining comparative subjective judgments of the noises. The advantage of such an experimental setup is that many of the physical factors in the tests are under very close control. The disadvantage, however, is that the environment lacks realism and thus may influence subjective judgment.

Another example is the field study which involves special flight operations arranged so that judgment data can be obtained. In one of these studies about a dozen different aircraft, including helicopters, were used, and subjects were located both outside and inside residence-type structures for subjective reaction studies.

These studies were more realistic for the subjects but obviously did not completely represent real-life situations. Current, and future, testing for subjective responses are making use of improved capabilities for realistically simulating both indoor and outdoor noise exposure situations. The NASA sponsored Noise Research Laboratory at Columbia University and the Aircraft Noise Reduction Laboratory now under construction at the Langley Research Center will provide a high degree of environmental control for this type study.

Studies of noise induced sleep interference use electroencephalographic (EEG) meas- urements as an indicator of sleep state. These studies correlate the EEG pattern shift with the subject's noise exposure history during sleep. Realistic laboratory environ- ments (Stanford Research Institute; North Carolina State University) are provided for the subjects who are exposed to various types and levels of aircraft noise.

Currently the laboratory data are being supplemented by a study that is being conducted in actual homes near airports (refs.

10 to 14).

MEASUREMENTS OF NOISINESS Three general types of noisiness measurement units for flyover noise exposures have been proposed. They are characterized as maximum units, effective units, and composite units. The factors considered in defining these noise evaluation units can be further categorized as seen in figure I-1.

The maximum units include such measurements as A-scale, which along with a number of other similar units can be measured by a meter having an appropriate filter system. The function of the filter is to provide frequency weighting to represent the noisiness value of the noise. These units are useful in providing a quick determination of characteristic psychological responses of people to noise. Figure I-2 illustrates the significance of A-scale levels in dB(A) in traffic noise situations. Figure I-3 on the other hand provides comparative perceived noise levels at various distances for several types of transportation vehicles.

The effective units, on the other hand, are designed especially for aircraft noise certification. They incorporate such additional features of a single noise exposure as its audible pure tone content, the duration of exposure, and other significant aspects of its time history (fig. I-3).

They are evaluated by a more sophisticated data analysis, which is usually performed by a computer.

An example of an effective unit is the effec- tive perceived noise level (EPNL). In studies to date, a number of effective units for noisiness measurements have been developed and evaluated.

Composite units are designed to represent the overall exposure, at a point or over an area, to series of noise events which occur during a given period of time.

Thus, additional adjustment factors accounting for the number of overflights, the types of air- craft involved, and their operating schedules are incorporated.

Computer programs are available to predict noise contours having equal values of such composite units as Composite Noise Rating (CNR) or Noise Exposure Forecast (NEF).

Noise contours are useful in determining the ground areas and associated populations that receive given noise exposures.

(See, for example, fig. I-4.) The ability to predict such noise con- tours can be an important consideration in land-use planning for airport communities.

Valid contour predictions infer an understanding of such atmospheric effects as atten- uation, scattering, and refraction, and the influence of terrain. Simplifying assump- tions are usually made regarding all the above factors, and as a result prediction accuracy is degraded as distance is increased.

In order to evaluate these measurement units, NASA has supported several labora- tory and flyover studies, which included some evaluations of the above units. For each of these studies, hundreds, or sometimes thousands, of judgments were made, and the results were evaluated against various measurement units. In these studies the effective units were generally the most accurate.

MAXIMUM EFFECTIVE COMPOSITE UNITS UNITS UNITS ABSOLUTE NOISE LEVELS PNL A - Scale NOISE SPECTRUM N - Scale AUDIBLE PURE TONES EPNL FLYOVER DURATION I TIME HISTORY AIRCRAFT TYPE CNR MIX OF AIRCRAFT NEF NUMBER OF OPERATIONS RUNWAY UTILIZATION FLIGHT PATH OPERATING PROCEDURES TIME OF DAY Figure I-1. - Characteristics of noise stimuli that are accounted for by various types of sub- jective evaluation units.

NOISE LEVEL, dB(A) COMMUNITY ACTIVITY AND LEGAL ACTION PETITION OF PROTEST 80- LETTERS OF PROTEST COMPLAINTS LIKELY POSSIBLE COMPLAINTS ----- - -- - -70- RARE COMPLAINTS 60 ACCEPTANCE 50- Figure I-2. - Estimated community responses for various levels of single event traffic noises (ref.

11).

o OUTDOOR PERCEIVED NOISE LEVEL, PNdB 50 - 500 1000 DISTANCE FROM VEHICLE, FT Figure I-3.

- Example perceived noise level as a function of distance from several transportation vehicles (ref. 9).

/N 30

SCALE: 1 IN. = 12000FT Figure I-4. - Noise exposure forecast (reference) contours for 1970 and 1975 operations at Raleigh-Durham, N.C., airport, reflecting effects of increases in traffic volume on ex- posed areas (ref. 8).

COMMUNITY REACTION TO AIRCRAFT OPERATIONS studies to determine NASA has sponsored a number of community survey research the manner in which people react to airport noise. This work has been accomplished through contract support of the National Opinion Research Center, the University of Chicago, Columbia University, and Tracor, Inc. In these studies, random sampling procedures have'been employed to evaluate the reactions of people in various kinds of environments for which the noise exposures could be characterized. To date, studies have been conducted in nine airport communities for which the environmental situations varied widely. Both large and small cities were included in the study as well as noisy and relatively quiet neighborhoods.

The reactions of individuals to aircraft noise exposures were noted to correlate with several factors other than their noise environment. These nonnoise factors (which were interrelated with noise level) included fear of aircraft crashes, noise susceptibility, noise adaptability, misfeasance, and pollution annoyance. For all cities studied, pertinent data relating to those highly annoyed and to complainants are listed in table I-1. The percent of persons highly annoyed was always greater than the percent who complained, as is indi- TABLE I-1. - ANNOYANCE AND COMPLAINT DATA DERIVED FROM INTERVIEWS IN AIRPORT COMMUNITIES OF NINE CITIES (REF. 8) City Number Number Percent Number Percent inter- highly highly com- com - viewed annoyed annoyed plainants plainants CHI 872 299 34 43 4.9 DAL 923 236 26 22 2.4 DEN 1009 215 21 33 3.3 LAX 786 382 49 93 11.8 BOS 1166 517 44 156 13.4 MIA 676 148 22 12 1.8 NYC 1070 696 65 240 22.4 CHAT 1114 102 9.2 3 0.3 RENO 846 124 14.6 4 0. 5 cated in figure I-5 which is plotted from the data of table I-1. Given the noise exposure (for CNR values between 85 and 125) and population density patterns for a community, the number of highly annoyed persons and complaints can be predicted (ref.

8). High annoyance was not detected in any of the test sites that had a CNR value of 85 or less.

The CNR values corresponding to a high annoyance level for 20 percent of the surveyed population ranged from about 98 for the large cities to 115 for the small cities.

oNEW YORK CITY LOS ANGELES o BOSTON PERCENTAGE 40 OF PEOPLE HIGHLY CHICAGO ANNOYED 30 o DALLAS o MIAMI 20- 0o DENVER RENO CHATTANOO GA I l I I I 0 5 10 15 20 25 PERCENTAGE OF COMPLAINANTS Figure I-5. - Relation between percentage of highly annoyed people and percentage of complainants in statistical samples of the population of nine cities.

SONIC BOOM The technology for sonic boom assessment has not been precisely developed; how- ever, effort has been expended to characterize the statistical nature of the exposure.

Important in this exposure is the variability from a true N-wave along with associated community and individual responses.

NASA personnel have had a monitoring, super- vising, or coordinating role in all major research projects relating to the effects of sonic booms on community response.

Other agencies which participated actively and provided substantial resources are the Department of Defense, Department of Trans- portation, and the Department of Commerce.

Laboratory simulation studies, jury studies involving people both inside and outside residential structures, and actual supersonic overflights of entire communities for the purpose of evaluating human responses have been performed.

The sonic boom is unique among noise exposures because it is impulsive in nature, exposes very large areas nearly simultaneously, and comes without warning.

It can have a direct effect on people who are exposed in an outdoor situation and a much differ- ent effect on those who are sheltered inside a building or other structure. Detrimental effects on buildings and on other possessions may in certain circumstances be very significant in shaping public opinion and in defining acceptance criteria. (See ref. 15.)

INDOOR LISTENING OUTDOOR LISTENING 100 - 8 XB-70 0 F-104

60- * B-58

40 - O ap - '60 PERCENTAGE OF PERSONS RATING SONIC BOOM 2 NOMINAL SONIC BOOM OVERPRESSURE, NIM UNACCEPTABLE 100- 50 100 200 0 50 100 MEDIAN OUTDOOR MEASURED SONIC BOOM OVERPRESSURE, N/M Figure 1-6.

- Percentage of people who rated as unacceptable sonic booms from XB-70, F-104, and B-58 aircraft. (Listeners from Edwards Air Force Base. ) All aircraft were the first aircraft of a pair. Data from reference 14.

Acceptability of sonic booms with respect to annoyance for both inside and outside listening situations was studied using actual aircraft overflights, and the results are summarized in figure 1-6. For a projected rate of 10 to 15 booms per day, the percent- age of unacceptability ranges from virtually zero at measured peak overpressure of 36 newtons per square meter (0. 75 lb/ft ) to virtually 100 percent at a peak overpressure of 172 newtons per square meter (3.6 lb/ft ).

The causes of annoyance among those expressing more than a little annoyance during community flyover tests are divided into categories in figure 1-7. Such factors as house rattles, startle, and interruption of sleep and rest are noted to be significant.

Studies of human response to sonic boom have identified the significant parameters of the boom pressure signature.

Outdoor annoyance increases markedly as the "risk time" of the shocks decreases as well as with the degree of "spikiness" in the signa- ture. Indoor annoyance, on the other hand, depends upon the loudness level, the degree of rattle and vibration and, among other things, whether the individual is a homeowner.

For all supersonic aircraft tested to date, neither rise time nor random spikiness NUMBER OF RESPONDENTS 336 686 1702 1376 I I I I 7 - 6C0- SALL CAUSES HOUSE RATTLES 5C - PERCENTAGEOF 40- RESPONDENTS REPORTING ANNOYANCE STARTLE 20- INTERRUPTS SLEEP 1 AND REST 10 - INTERRUPTS CONVERSATION INTERRUPTS RADIO, IV 0 I I I I .6 .8 1.0 1.2 1.4 1.6 LBI FT MEDIAN PEAK OVERPRESSURE, I I I I I I 30 40 50 60 70 80 MEDIAN PEAK OVERPRESSURE, N/M Figure I-7. - Factors contributing to annoyance due to various levels of sonic boom exposure during community overflights (ref. 14).

in the signature are controllable features; they depend on random atmospheric effects of turbulence, wind, and temperature. The received signature depends also on further random factors such as the configuration of buildings and terrain and whether the boom is heard outdoors or indoors.

For a single flight, the only feature subject to a degree of control is the flight profile and, therefore, the nominal peak overpressure (and other parameters) that may be calculated as characteristic on the average of that flight profile.

Human response measured against nominal mean overpressure obtained from community surveys is therefore the most pertinent indicator of public response to the sonic boom at a given boom frequency.

The present data on community response to the sonic boom lack adequate informa- tion pertaining to boom frequency and to nighttime sonic boom exposure as well. The data presented in figures I-6 and I-7 provide a guide to acceptability.

RESEARCH PLANS LABORATORY STUDIES Because of current design trends, it is expected that future configurations of air- craft will generate noises that are closer in level to the background noise in the com- munity and will have noise spectra with relatively more intense low-frequency noise components than do current aircraft.

Laboratory studies are planned, with the use of improved facilities, including those of the Langley Aircraft Noise Reduction Laboratory, to study the intrusiveness of aircraft noise, in particular the significance of background noise and the effects of low-frequency noise and noise induced vibrations on the psycho- physiological responses of people.

By means of flyover noise simulation techniques, tests are being conducted to evaluate proposed retrofit modifications to existing aircraft as well as proposed new configurations and to establish a close correlation between the reactions of people in laboratory simulation situations and real life exposures.

AIRPORT COMMUNITY NOISE STUDIES Both short range and long range plans have been developed for airport community noise research. Data will be obtained by means of special tower facilities to define more precisely the propagation losses and the variability of noise propagating through an inhomogenous medium from flight altitudes to the ground at various angles. These data will be correlated with actual ground contour measurements from aircraft in flight in order to improve the capability for predicting contour patterns, particularly at large distances. Long range plans call for repeating community surveys in selected localities in order to evaluate and correlate expected changes in the noise exposures and the associated responses.

Preliminary tests are showing that there are positive benefits in retrofit as depicted by subjective rating scales. Unpublished data have shown that a 6 EPNdB retrofit reduction from a base noise level is perceived as less annoying at the 1. 1, 2. 5, and 3.

5 mile measuring points on approach path. Further studies will be conducted with other aircraft, including their various takeoff and landing noise spectra.

The aspect of fear is also being studied to determine its role in judgments of aircraft annoyance.

SONIC BOOM STUDIES Most of the sonic-boom-related human response studies to date have been focused on the practical operating range of the current generation of supersonic transports for which the associated sonic boom intensities are in the range 1. 5 to 4.0 pounds per square to evaluating the responses of people to foot. Very little effort to date has been applied overpressures that may be achievable with the application of new design and sonic boom construction technologies. Planned research will include studies of a wide range of to sonic boom intensities.

human responses REFERENCES 1. Kryter, Karl D.: The Effects of Noise on Man. Academic Press, Inc., 1970.

S.; Horonjeff, Richard D.; and Bishop, Dwight E.: The Noisiness 2. Pearsons, Karl of Tones Plus Noise. NASA CR-1117, 1968.

Effects of Tone and Duration Parameters on 3. Pearsons, Karl S.: Combination Perceived Noisiness. NASA CR-1283, 1969.

4. Kryter, K. D.; Johnson, P. J.; and Young, J. R.: Judgment Tests of Flyover Noise from Various Aircraft. NASA CR-1635, 1970.

5. Hinterkeuser, Ernest G.; and Sternfeld, Harry, Jr.: Subjective Response to Synthesized Flight Noise Signatures of Several Types of V/STOL Aircraft.

NASA CR-1118, 1968.

6. Young, J. R.: Attenuation of Aircraft Noise by Wood-Sided and Brick-Veneered Frame Houses. NASA CR-1637, 1970.

7. Kryter, K. D.: Possible Modifications to the Calculation of Perceived Noisiness.

NASA CR-1636, 1970.

8. Connor, W. K.; and Patterson, H.: Community Reaction to Aircraft Noise Around Smaller City Airports. NASA CR-2104, August 1972.

9. Wilson, A. H.: Noise, Final Report of the Committee on the Problem of Noise, London, 1963.

10. Lukas, Jerome S.; and Kryter, Karl D.: A Preliminary Study of the Awakening and Startle Effects of Simulated Sonic Booms. NASA CR-1193, September 1968.

11. Lukas, Jerome S.; and Kryter, Karl D.: Awakening Effects of Simulated Sonic Booms and Subsonic Aircraft Noise on Six Subjects, 7 to 72 Years of Age.

NASA CR-1599, May 1970.

12. Lukas, Jerome S.; Dobbs, Mary E.; and Kryter, Karl D.: Disturbance of Human Sleep by Subsonic Jet Aircraft Noise and Simulated Sonic Booms. NASA CR-1780, 1971.

July 13. Lukas, J. S.; and Dobbs, M. E.: Effects of Aircraft Noises on the Sleep of Women.

NASA CR-2041, June 1972.

14. LeVere, T. E.; Bartus, Raymond T.; and Hart, F. D.: The Relation Between Time of Presentation and the Sleep Disturbing Effects of Nocturnally Occurring Jet Aircraft Flyovers. NASA CR-2036, May 1972.

15.

Ribner, H. S., et al.: Sonic Boom. Second Report of the Sonic Boom Panel, ICAO, Montreal, 1970.

II - SOURCE ABATEMENT TECHNOLOGY

SUBSONIC TRANSPORT AIRCRAFT (and for the next decade) is generated by the The noise we hear and object to today aircraft. Accord- subsonic conventional takeoff and landing (CTOL) commercial fleet of Engine and Refan discussion on CTOL noise in reference to the NASA Quiet ingly, the Programs is followed by a discussion of noise associated with supersonic aircraft, Much of the work discussed under short haul powered lift aircraft, and rotorcraft.

application and will lay a gen- subsonic aircraft is germaine to these other areas of noise contributors; for example, the eral groundwork with regard to certain common Subsonic aircraft by fans and jets and methods of sound absorption.

noise generated The in terms of propulsion system noise and airframe noise.

noise will be considered in-house activities material covered herein reflects a total NASA posture, including both work done under contract.

and PROPULSION SYSTEM NOISE the engines has been and continues to be the major source of air- The noise due to in figure craft noise. The sources of noise in a typical turbofan engine are indicated also the compressor blades propagates for- II-1. Noise created by the fan blades and is radiated to the community. Fan noise also prop- ward to the cowl inlet and thereafter the fan exit nozzle and is radiated. In addition, the high velocity jet agates rearward to Noises from of air from the fan mixes violently with the ambient air and creates noise.

the combustor and turbine propagate rearward to the core nozzle exit and then are radi- to the community. The very high velocity of the core gas jet produces one of the ated (not shown) are radiated outward through the en- major sources of noise. Other noises created by the fan discharge air flowing at gine and nacelle structure. Finally, noise is con- high velocity over the aft surface of the nacelle afterbody. The material to follow siders the various noise sources, what is being done and is planned to reduce the noise generated from these sources, and techniques used to absorb noise which cannot be Consideration will also be given to the performance or eco- eliminated at the source.

employed to reduce these noises.

nomic penalties incurred by the measures JET NOISE r FAN INLET FAN DISCHARGE .

CORE J NOISE FAN NOISE NOISE TURBINENOISE-' NOISE COMPRESSOR CS-64045 noise sources.

II-L - Turbofan Figure SUBSONIC BLADE PASSAGE FREQUENCY I 15dB SOUND FAN BROADBAND NOI E PRESSURE LEVEL, SUPERSONIC - MULTIPLE PURE TONES dB -BLADE PASSAGE FREQUENCY NOISE 8,,.ROAD BAN D LFA.N 0 2 4 6 8 10 FREQUENCY, kHz CS-63223 Figure II-2. - Fan noise spectra at subsonic and supersonic tip speeds.

Fan Noise The first and most important function of a fan is to produce pressure rise. Our concern is how to do this noiselessly and without undue penalty. In a fan designed to produce a given pressure rise, there are two parameters that may be traded. These are the fan tip speed and the overall stage loading, conveniently measured by a work co- efficient given by the ratio of the average turning of the airstream to the fan tip speed.

Early studies of fan noise indicated that tip speed was highly important in determin- ing the noise.

Figure II-2 shows spectra for two fans, one having a subsonic tip speed and one having a supersonic tip speed. Considering first the subsonic tip speed fan, it can be seen that the noise energy appears as discrete tones that are superimposed on a broadband noise.

The discrete tones are the cause of the characteristic whine of cur- rent aircraft heard on landing. These same noises are seen in the spectra from the supersonic tip speed fan, but a large number of additional discrete tones or spikes are also evident. These additional tones are associated with the shock waves on the blade leading edges and occur at multiples of the rotational speed of the fan.

They may be heard on wide-bodied aircraft at takeoff and are characteristically called buzz-saw or multiple pure tones.

It can be seen from these spectra that tip speed is indeed a factor in fan noise. In fact, two noise regimes exist depending on whether the fan operates at subsonic or supersonic tip speeds.

Several subsonic speed fan stages have been noise tested at NASA along with one supersonic tip speed fan. The data from these fans provide a good indication of the state-of-the-art in fan noise.

It is necessary to consider the aerodynamic and noise features of these fans as these features qualify the data.

First, it should be noted that the fans were all 6 feet (ft) in diameter.

This means that no scaling of the data is required.

They are all designed without inlet guide vanes since the extra blade row was a potential noise source.

Rotor-stator axial spacing was at least two rotor blade chords to allow rotor wakes to decay before impinging upon the stators. While this wide spacing reduces noise about 3 decibels (dB), it also serves to add length and weight to an engine. There also is some indication that this spacing costs about 1 percent in fan stage efficiency.

Finally, the ratio of stator vane to rotor blade number was selected on most of the fan stages so that blade-passage frequency tones were cut off (do not propagate to the far field) according to the Tyler-Sofrin theory (ref. 1). Cutoff requires that the vane blade ratio be about 2. This requirement leads to thin, narrow stator vanes that create structural problems and affect aerodynamic per- formance. In one case, stall margin was reduced from 16 to 7 percent by meeting the cutoff vane number requirement. The aerodynamic design of the fan stages were state- of-the-art.

Specific flow rates were generally about 41 pounds per square foot per sec- ond (lb/(ft )(sec)). Local values of the diffusion factor were not allowed to exceed about 0.5.

Figure II-3 shows the matrix of fans tested in terms of the overall performance pa- rameters. Fan stage pressure rise is plotted against tip speed with the average work coefficient shown as a parameter.

It can be seen that a range of pressure ratios (work coefficient) is covered at constant tip speed and that a range of tip speeds (work coeffi- cient) is covered at constant pressure ratio. Several of the fans also lie close to a line of constant work coefficient.

Figure II-4 shows the maximum perceived noise level along a 1000-ft sideline as a function of fan pressure rise. These data have been normalized to unit thrust. The correlation for subsonic tip speed fans shows a total spread of about ±2.5 perceived noise decibels (PNdB).

This correlation shows that perceived noise increases by 3 PNdB for each doubling of the thrust and by 4.2 PNdB for each doubling of fan pres- sure rise. An equivalent functional dependence can be obtained in terms of fan diameter and tip speed.

In these terms, the noise increases by 15 PNdB for each doubling of tip 0.6/.5.4 .3 WORK - COEFF, .6 - o FAN A (V 6 o FAN B * FAN C TOTAL .4 - a QF-1 PRESSURE , QF-3 RISE, , QF-5 PR - 1 - QF-6 > QF-9 .2 I I I I I l l iIII .1 I 1000 1500 2000 SPEED, Ut, FTISEC TIP Figure II-3. - Fan design parameters.

UT, FTISEC o FAN A 1160 125- o FAN B 1160 o FAN C 1550 6 QF-1 1107 0o9/ 5dB 120- q QF-3 1107 / 0 QF-5 1090 o t QF-6 / 5 > QF-9 700 115 - SOLID SYMBOLS DENOTE / / SPEED/ 90% DESIGN 10 log F PWL - 110 - PWL = 121.9 + 14 log(P - 1) + 10 log F 105 - R I I I ll I I I I I 1.0 .01 .1 FAN PRESSURE RISE, PR - 1 of perceived noise from low tip speed fans.

Figure II-4. - Correlation Maximum value, 1000 ft sideline.

speed and 6 PNdB for each doubling of diameter. These fans, designed for low noise, have demonstrated the lowest noise levels known.

As fan tip speed moves into the supersonic range, a new noise source emerges. As noted earlier, this source is associated with the shocks on blade leading edges. Because of the blade-to-blade manufacturing differences and localized flow differences, the shocks are not identical with the result that some overtake'and merge with others. This leads to the multiple-pure-tone signature observed in the far field.

Current fan design procedures require supersonic tip speeds for single-stage fans at fan pressure ratios above about 1. 6, and they may be employed at pressure ratios as low as 1.4. The upper limit on pressure ratio with current design techniques is thought to be about 1.9. At larger pressure ratios, multistage fans are required. This stage addition allows a reduction in tip speed relative to that required for a single-stage fan at the same pressure ratio.

The result is that the reduced tip speed may offset the noise increase caused by the addition of a second stage.

There are no data available on low- noise, two-stage fans to permit this comparison. In fact, an objective of the future pro- gram is to obtain the comparison.

The two-stage fan data shown were obtained from older fans, such as those on the JT3D and JT8D engines.

Figure II-5 includes perceived noise from one-stage supersonic tip speed and two-stage fans as functions of fan pres- sure ratio. The data are at the maximum 1000-ft sideline and for a 90 000-lb thrust.

For completeness, the subsonic-tip-speed data in figure II-4 are also shown at the fan tip speed corresponding to takeoff operation.

The noise increment in passing from single-stage, low-tip-speed fans to single- stage, high-tip-speed fans is reasonably well documented.

One fortunate circumstance is that the multiple-pure-tone noise causing the high-tip-speed-fan-noise increment is very amenable to suppression by acoustic treatment.

A relatively small amount of treatment is effective.

The noise level increment incurred with two-stage fans is de- rived from data for fans with few or no noise reduction features in their design. Studies indicate that these levels may in fact be no more than for the high-tip-speed, single- stage fans.

Future work is intended to resolve this point.

Figure II-5 shows that for a 90 000-lb thrust a variety of fans can be selected to produce noise levels at a 1000-ft flyover in the 100 to 120 PNdB range.

Selection of a particular fan depends on the fan configuration, for example, stage number, tip speed, and the design pressure ratio, required.

The mission to be performed will exert a large influence on the fan selected.

In the Quiet Engine Program, NASA has examined several techniques for reducing fan noise. These have included rotor leading edge serrations, casing boundary layer 120 -oTWO-STAGE 1SINGlE-STAGE HIGH-SPEED MAXIMUM NOISE 110 SINGLE-STAGE PEEB 10 LOW- SPEED PNdB 100 - 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 FAN PRESSURE RATIO CS-63398 Figure II-5. - Fan noise levels. Thrust, 9000 Ib; ft flyover.

bleed, and leaned stators. Of these, only leaned stators have thus far shown any sig- nificant overall noise reduction. Even here the reduction was only of the order of 2 dB.

However, leaning the stators did result in one to two points of improvement in fan effi- ciency, and this is of interest. All of these techniques and others need to be further ex- plored in detail to determine designs that will yield maximum noise benefits.

Future efforts are in this vein. Industry and university responses are anticipated currently on basic research efforts being solicited in several noise problem areas in- cluding fan noise. Additionally, a contracted effort is anticipated shortly on new and novel techniques for fan source noise reduction.

Jet Noise Jet noise can be one of the major contributors to the total noise generated by sub- sonic CTOL aircraft operating near airports. The level of jet noise is primarily deter- mined by the jet exhaust velocity (fig. II-6). Consequently, the high exhaust velocities of the older CTOL aircraft and SST engines make it difficult for those engines to meet current FAR-36 standards. On the other hand, advanced CTOL engines have lower ex- haust velocities so that the noise goals can be more readily met so far as jet noise is concerned.

It should also be mentioned at this point that the forward speed of the aircraft lowers the jet velocity relative to the airstream and therefore reduces the noise level. Some effects of forward speed on noise generation are shown in figure 11-7, which compares the noise spectrum from a 2-inch (in.) circular nozzle with an airspeed of 100 knots (KTS) to the case for no forward speed. Forward speed was simulated by placing the 2-in. nozzle in a large free jet (ref. 2). A forward speed comparable to takeoff (100 KTS) clearly reduces the high frequency jet noise generation. Experiments on the effect of forward speed are continuing at NASA and elsewhere.

In addition to the level of the noise the frequency of the noise is important. Noise near a frequency of 3000 hertz (Hz) sounds loudest to the human ear. It sounds decreas- ingly loud at lower and higher frequencies. Therefore, noise annoyance is primarily reduced by reducing the noise level; but it can also be reduced by keeping the frequency of the noise far away from frequencies having high annoyance weighting.

The general methods of reducing subsonic jet noise level when necessary are: (1) Reduce the noise generated (2) Redirect the noise already generated away from the community (3) Absorb the noise already generated The specific problem areas and the progress made in understanding subsonic jet noise, and in particular its suppression, are now discussed.

Figure II-8 contains sketches of common nozzle shapes used on engines, namely the 120 - 110 - JET NOISE, 100 - CTOL PNdB 6 AIRPLANES BYPASS I I I RATIO I 600 800 1000 1200 1400 1600 cs-s59064 EXHAUST JET VELOCITY, FTISEC Figure II-6. - Effect of engine exhaust velocity on jet noise for a 500 ft flyover.

Airplane gross weight, 100 000 lb. Total thrust (4-engines), 60 000 lb.

100 - 90 - SOUND * .

PRESSURE A A AA *A LEVEL , AA dB AIRSPEED, A KTS 70 - 0 ^ 100 60o L I I I 10000 FREQUENCY, Hz Figure II-7. - Effect of airspeed on noise spectra of a model, 2 in. diameter convergent nozzle. Exhaust velocity, 835 ft/sec.

Angle from inlet, 1000.

CIRCULAR -SLOT PLUG cs-6s5600oo Figure II-8.

- Common types of nozzles.

160- = THEORY(n= -3, i 0.62)-\ 150- o 140 -' TOTAL SOUND POWER LEVEl.,IN POWER LEVEL, / NOZZLE DIAMETER, IN.

dB 130- 0o 1.06 a 1.625 a 2.06 o 3.06 L 4.0 v 13.0 + .98 (LUSH) 110.

I I I I CS-697 400 600 800 1000 1200 CS-65597 NOZZLE EXHAUST VELOCITY, FTISEC Figure II-9.

- Comparison of jet noise power for circular, slot and plug nozzles. (All data scaled to an area of 1 ft and 770 F.)

o GRADUAL INLET, THIN LIP, C . 0.98 STANDARD SHAPED v 0 GRADUAL INLET, THICKER LIP, C 0.98 NOZZLES v o GRADUAL INLET, VERY THICK LIP, Cv 0.98 o ABRUPT ROUND INLET, C - 0.92 v 120 - CONE INLET, SHARP LIP, C . 0.85 v - -- - SHARP EDGED ORIFICE, C * 0.68 v - 1 \ dB SOUND POWER LEVEL, 100 - dB C R 1 C R C 2 R 2 3 3 80 I I III I I .1 .2 .4 1 2 4 10 20 cs-65605 kHz FREQUENCY, Figure II-10.

- Effect of nozzle inlet and lip shape on noise.

Nozzle exhaust velocity, 785 ft/sec; diam, 15/8 in.

circular, slot, and plug nozzles. The first question to be answered is whether any of data for the total these generate less noise. This is answered in figure II-9. Typical noise power generated by each of these nozzle shapes are plotted as functions of the jet exhaust velocity. Only a very small sample of the large amount of subsonic ambient temperature jet noise data, taken at the NASA Lewis Research Center and elsewhere, has been plotted here. The equivalent area diameter of the nozzles ranged from about 1 in. to 1 ft; therefore, the data have been normalized to a common area of 1 ft . It is clear from this figure that the total noise power generated at a given subsonic velocity by slot and plug nozzles is the same as that generated by circular nozzles of the same area. A circular nozzle typically has somewhat better aerodynamic performance.

Thus, it is doubtful that a reduction in jet noise generation can be achieved by using either slot or plug nozzles.

The nozzles described in figure II-8 are of good aerodynamic design in that they have a gradual area contraction and a thin lip. To aid our understanding of jet noise it was helpful to see what effect a thick lip and a nongradual inlet would have on the jet noise generation (ref. 3). Figure II-10 contains plots of the noise power spectra for a number of circular nozzles of the same diameter at the same peak or isentropic jet ve- locity. The data are for nozzles that differ only in their inlet shape and lip thickness.

Except for the sharp edged orifice all generate the same noise. This similarity also oc- curs for the noise radiation patterns and at higher and lower subsonic velocities than plotted in figure II-10. This conclusion can probably be carried over to slot and plug nozzles.

Although the example nozzles of figure II-8 show little difference in noise generation characteristics, there are some other nozzle types that appear to generate less noise.

For instance, multitube nozzles are known to be useful for noise reduction. Evaluations of such nozzles have been accomplished on an ad hoc basis and verifications and improve- ments have been made by parametric variations. It has been shown that multitube sup- pressors shift the frequency of the spectrum peak upward compared to that of the basic nozzle as shown in figure II-11. This frequency shift is beneficial because of the asso- ciated increase in atmospheric propagation losses and the relatively higher attenuation provided by sound insulating structures. There is also the potential for optimizing the spectra for subjective reaction purposes. The mechanisms of noise suppression are still not well understood; however, it is found that the noise sources are concentrated closer to the multielement nozzle exit than for an equivalent single element nozzle exit.

This suggests that lined ejectors will probably be more effective in reducing noise for the multitube nozzles than for the standard nozzle.

Another suppressor nozzle is the coannular nozzle. Coannular nozzles are quieter than two separate jets having the same total throat area. The noise level is dependent on the fan to core area ratio and on the fan to core velocity ratio. Nozzles with large fan to core area ratios are quieter. The coannular nozzle is quietest when the fan to 10 - ,,SINGLE ROUND NOZZLE ACOUSTIC 10-6 POWER SPECTRAL UBE SUPPRE 21-TUBE SUPPRESSOR DENSITY, NOZZLE W/Hz -7 10- 103 104 105 FREQUENCY, Hz Figure II-L - Effects of nozzle geometry on radiated acoustic power of constant flow momentum and constant area. Area, 4. 52 in. 2; jet velocity, 694 ft/sec.

core velocity ratio is about 0.5.

Unfortunately it is usually not practical to use an en- gine cycle with such a low velocity ratio (0.7 to 0.9 is more common).

Another method of alleviating jet noise is to redirect some of the energy away from the observer on the ground.

This can be accomplished by means of ejector shields, by rigid surface shielding such as that provided by a wing (ref. 4), and by means of reflec- tions due to impedance discontinuities in the flow.

Experimental and theoretical work is presently being conducted relative to the re- direction of jet noise, and all three of the previous approaches are being studied.

Par- ticular attention is being given to experimental evaluation of reflections caused by a sound gradient interface between the jet and the observation points.

Laboratory experi- ments using helium to simulate an impedance discontinuity in a jet have produced high transmission losses over a broad range of frequencies.

Satisfactory results have been obtained using both point sources and a model scale jet.

Experiments which have pro- duced useful noise reductions due to wing shielding of the noise from top-of-the-wing mounted jet exhausts for both subsonic and supersonic jet flow velocities are continuing in scale model studies.

The third method to reduce the jet noise is to absorb some of the generated noise.

This can be done by exploiting the absorption of high frequency noise by the atmosphere and by employing acoustic linings.

For example, the previously mentioned ejector shield can be lined with sound absorbing material.

Furthermore, attempts can be made to bend the sound rays (by velocity or temperature gradients) toward sound absorbing material or so that they take a long path through the atmosphere to increase the noise reduction.

In summary, subsonic jet noise can be reduced by employing high bypass ratio en- gines with low engine exhaust velocities. With existing engines the jet noise can be re- duced somewhat by techniques which reduce the generation of noise and/or redirect and absorb it.

NASA is currently conducting in-house research to develop better methods and tech- niques of suppressing low velocity jet noise.

Other Noise Sources Fan and jet noises have now been reduced in research engines to the point that other noise sources, such as turbine and combustor noise, become evident.

In fact, with quiet engine A, a 6300-Hz turbine tone stands out of the overall noise spectrum when the engine is operating in the fully suppressed nacelle. The turbine tone contributes mea- surably to the perceived noise at the rear of the engine. These noise sources are now starting to be investigated in a systematic manner.

The discussion to follow is based primarily on what has been gleaned from analysis of complete engine data.

Both discrete tone and broadband noise emissions have been observed originating within the engine core. The discrete tones originate with the rotating machinery, and broadband noise can also be generated. In addition, broadband noise can be generated by combustion processes and by obstructions and discontinuities in the flow path.

The existence of these other noise sources has been demonstrated with data of the kind shown in figure II-12, which shows how engine noise changes with jet velocity.

For many years it was widely held that all the noise measured was jet noise, and that at low velocities the noise changed only slightly with changes in velocity, while at higher veloc- ities the change in noise was greater.

Recent work at NASA has shown that, except for very high speed jets, jet noise per se follows an eighth power relationship at all speeds, which is consistent with theoretical arguments.

The higher than predicted noise at low velocities, which has been reported as jet noise by many investigators, is in fact of core-engine origin and reflects a faulty interpretation of the experimental data.

More detailed spectral analyses of such data confirm these findings.

A typical engine sound spectrum appears as figure II-13. It is usually dominated by fan machinery and core jet noise as illustrated; but, other spectral components, tones in particular, can be shown to exist and are identified with the blades of the turbine.

For engines where the fan and jet noises have been reduced, these turbine-associated noises play a significant role in annoyance such that full potential of further reductions of fan and jet noises cannot be realized unless the turbine noise is also reduced.

0 MEASURED ENGINE NOISE - -- THEORY AND RECENT TESTS

SOUND _F

POWER 10 dB LEVEL

owp

"1

I

, 500 700 1000 1500 2000 JET VELOCITY, FT/SEC Figure 11-12. - Jet noise.

JET Fr JET -CORE SOUND /-TURBINE PRESSURE LEVEL BYPASS JET 10 000 FREQUENCY, Hz Figure 11-13.

- Spectral components.

If this turbine noise has been correctly identified, two techniques used to reduce fan noise may be applicable in reducing turbine noise. The first involves reducing the tones by spacing out the adjacent blade rows in the turbine by mechanical design. The second method employs the application of sound absorbing materials, or core suppressors, to the engine. Wide spacing of turbine blade rows is costly to demonstrate and involves significant engine weight penalties in actual use. One experiment run under contract for FAA has demonstrated that significant turbine noise reductions are possible by wide spacing in a single-stage turbine (ref.

5) but also that the turbine efficiency was reduced significantly.

Some experiments have been conducted using sound absorption materials.

Figure II-14 shows the results of applying core suppression materials on a JT3D engine at ap- proach power.

An effective reduction of turbine noise is shown. A similar experiment was performed with quiet engine C. These data are shown in figure II-15 again demon- strating the potential of core suppression.

O UNSUPPRESSED * SUPPRESSED SOUND PRESSURE LEVEL 10 dB S m I I II I I II 1000 10000 FREQUENCY, Hz Figure II-14, - JT3D turbine noise suppression.

-_- o CORE UNSUPPRESSED SOUND PRESSURE LEVEL

II I

I I

500 100 500 1000 5000 10 000 FREQUENCY, Hz Figure II-15. - Quiet engine C sound spectra.

These preliminary results, though encouraging, address only one of many other po- tential sources.

Knowledge of the core noise problem is far from adequate.

The data of figure II-16 are presented as an example.

This shows the effect of the addition of a nacelle suppressor and a core suppressor to quiet engine A. Significant high frequency noise reductions are observed, but there remains a haystack in the spectrum whose ori- gin remains unclear. Other such anomalous spectral components have been variously observed and remain to be explained.

More can be expected to be discovered as testing progresses with the NASA quiet engines, but such tests offer the potential of identifying and resolving many of these problems also. Tests are planned for quiet engine C, for example, using a carefully designed core suppressor to pursue further the core noise problem.

The second noise associated with the core engine occurs at low exhaust velocities, influences the low frequency portion of the spectra, and is associated with internal con- ditions in the engine.

The fact that internal conditions can affect the low frequency noise has been demonstrated in several recent tests. Figure II-17 (from ref.

6) shows the be- havior of observed noise as a function of jet velocity for various internal flow conditions.

It shows noise levels which are controlled by the low frequency part of the spectrum.

Curve C was obtained from nozzle experiments in which considerable effort was ex- pended to eliminate upstream turbulence and flow noise. Curve B comes from tests of a simulated engine where the combustor cans were removed from the upstream piping.

Curve A represents data from a variety of engine and model rig tests. It is apparent that internal flow conditions have an effect on far field, low frequency noise.

Few tests have been performed to examine the effect of internal acoustic suppres- sion treatment on low frequency noise; however, test results from the NASA quiet fan (ref. 7) have shown that significant reductions of low frequency noise do occur as a re- sult of internal acoustic suppression.

This result was accomplished by the broadbanded nature of the suppressor in spite of the fact that it was designed for attenuation of high frequency noise.

A test to examine the effect of acoustic suppressors designed for low frequency at- tenuation was performed on a J65 turbojet engine (ref. 8). The results of the test are shown in figure II-18. A significant reduction of the low frequency noise is seen in the figure.

These experiences show that a portion of the low frequency noise, which has in the past been called jet noise, can be reduced or eliminated by appropriate internal modifi- cations. Some modifications which have been effective are streamlining flow passages and struts, as well as placing acoustic suppressors in the ducts for the purpose of at- tenuating low frequency noise.

A source of low frequency "core noise, "which can become important when the en- gines are operated at part throttle conditions during approach to an airport, is the noise generated in the combustor section of the engines.

0 UNSUPPRESSED SUPPRESSED * SPL S10 dB

I

I I

i I

500 1000 5000 10 000 FREQUENCY, Hz Figure 11-16. - Quiet engine A sound spectra.

A SOUND PRESSURE B LEVEL, dB C JET VELOCITY Figure II-17. - Low frequency noise as function of jet velocity (ref. 6).

O WITH

SUPPRESSION

o NO SUPPRESSION

SOUND PRESSURE LEVEL I.

I 100 200 FREQUENCY, Hz Figure II-1&8. - Spectra of noise showing effect of suppression on J65 engine (ref. 8).

Recent tests reported by Boeing in reference 9 show that burners (whether or not they are lit) contribute to the noise field. Typical Boeing data are given in figure II-19 for a large burner, a small burner, and no burner in the jet supply duct. The noise spectra at 25 ft from the exhaust nozzle is given with the flame off in II-19(a) and with the burners lit in II-19(b). The presence of the burner in the duct generates pressure drop (or turbulent eddy) noise even with,no flame - especially for the high resistance small burner. With the burners lit there is a large increase in low frequency noise 3.22 DIAM. NOZZLE Vj - 824 fps - LARGE BURNER 100 0 PR = 1.5, TT - AMB, o PR 1.5, TT= AMB, VJ= 824 fps -NO BURNER a PR 1.5, TT AMB, Vj = 824 fps - SMALL BURNER = 90- so - (a) FLAME OFF.

SPL(1) AT 3.22DIAM. NOZZLE T o * PR- 1.2, T - 935 F, Vi 930 fps -LARGE BURNER 0 F, 100 - o PR 1.2, TT 935 VJ 930 ALL BURNER = fp s - S M O PR 1.69, TT= AMB, Vj - 930 fps -NO BURNER 90 - 80q 70 - 60 I I I Ii II .250 .400 .630 1.0 1.6 2.5 4.0 6.3 10.0 16.0 25.0 40.0 FREQUENCY, f, kHz (b) FLAME ON.

Figure II-19.

- Burner noise.

caused by the presence of the flame accompanied by some reduction of high frequency noise.

NASA is initiating plans to conduct in-house combustor noise tests using existing facilities in order to determine means for predicting core noise levels and to find viable means of reducing the core noise floor. In addition, consideration is being given to an evaluation of industry experience in this area of jet noise and a compiling of available acoustic data. A current contract with Princeton University is also providing insight into the basic principles and problems underlying combustion noise (ref. 6). Noise gen- eration by a subsonic flow discharging from a combustion chamber was examined with regard to the relative importance of combustion as a source of noise in such a flow sys- tem. Measurements of pressure fluctuations inside the combustor were compared with far field noise measurements by direct cross correlations.

The cross correlations and derived cross-spectral densities verify that much of the noise originates inside the com- bustor. A first-order fluid mechanic perturbation model is used to predict exit plane velocity fluctuations due to internal pressure fluctuations.

This unsteadiness at the exit plane is assumed to behave as an acoustic monopole which radiates to the far field. Far field noise levels estimated on this basis are in good agreement with measured values.

The overall noise level from the combustor jet is found to be 10 to 20 dB higher than for an equivalent clean, cold jet at the same exit velocity.

As was mentioned earlier, several engines, such as quiet engine A, have been sup- pressed to the point that core engine noise, particularly turbine noise and low frequency noise associated with the various internal obstructions (combustors, struts, etc.), are limiting or determining the far field noise. Research and technology studies in these areas are only getting started. It is evident that cleaning the internal flow passages and the use of core suppressors can be effective in reducing these noises.

In thinking of the penalties associated with the reduction of these noises, it seems that core suppressors should not entail large penalties because they are involved with about 20 percent of the flow.

In fact, cleaning the flow passages should decrease losses and perhaps compen- sate for the suppressor.

Thrust reversing is used to shorten the landing distance for both conventional and short-haul powered-lift aircraft.

In addition, augmentor-wing STOL airplanes may use core-jet reversal to steepen the approach flight path. In the interests of minimizing the noise associated with aircraft operations, all potential noise sources should be con- sidered. Until recently, thrust reverser noise has received little attention.

NASA has initiated studies of thrust reverser noise (ref. 11).

Target-type rever- sers were used in the earliest tests because of their simplicity and because they can re- verse both circular nozzle (refs. 12 and II-20) and slot nozzle flows. Shown in figure II-20 are two reverser types that have been tested, namely, a V-gutter target and a semicylindrical target.

The noise directivity of a semicylindrical reverser is shown in figure II-21 as a SEMI-CYUNDER V-GUTTER CS-63253 Figure II-20. -Target-type thrust reversers.

OVERALL SPL REVERSED ----- (10-FT RADIUS), FLOW NOZZLE dB 00 t ! FLOW 1 o 110 120180 1001 90 100 120 110- NOZZIE ONLY/ ANGLE FROM NOZZLE AXIS,

K

CS-63236 directivity.

Figure II-21. - Thrust reversal noise Nozzle, 2 in.; jet velocity, 960 ft/sec; cylindri- cal reverser.

110- CYLINDRICAL REVERSER T 100 - SOUND PRESSURE LEVEL 90 - - AT 10 FT, p.c ' NOZZLE ONLY dB I I I I I I 70 (' 200 500 1000 2000 5000 10000 20000 THIRD-OCTAVE-BAND CENTER FREQUENCY, Hz CS-63244 Figure II-22. - SPL spectra for reverser and nozzle at angle of maximum sideline noise.

polar plot of the overall sound pressure level (OASPL) as a function of the angular posi- tion. The jet noise from the 2-in. circular nozzle alone has a pronounced directivity.

The maximum OASPL is 107 dB (re 20 N/m ) at an angle of 1600.

The minimum, toward the upstream direction, is 12 dB less than the maximum.

The reverser noise, in con- trast, is nearly uniform in all directions and is everywhere louder than the bare jet max- imum by 1 to 6 dB.

Toward the upstream direction, the noise level for the reverser is about 17 dB greater.

Maximum sideline noise spectra for the nozzle and reversed jets are shown in fig- ure II-22.

The sound pressure levels (SPL) for the cylindrical reverser exceed those for the nozzle over a broad frequency range, and its spectrum peaks at frequencies higher than those of the bare nozzle.

This peaking at higher frequencies increases the perceived noise level, but the higher frequencies attenuate more rapidly in the atmo- sphere than the lower frequency nozzle noise. Therefore, at short distances from the aircraft, such as the 500-ft sideline often used in defining STOL noise objectives, thrust reversal may well be a dominant noise source, while at distances appropriate for CTOL aircraft certification, atmospheric attentuation may reduce the reverser noise problem somewhat.

From the preceding information, it is obvious that thrust reversers can generate more noise than nozzles, at higher frequencies, and direct the noise more strongly toward the critical sideline and flyover points.

To present this problem in more specific terms, recent preliminary data on cascade reversers have been scaled up to the case of a CTOL aircraft of 300 000 lb gross weight using fan jet reversers on four NASA quiet engines.

Figure II-23 shows the perceived noise level (PNL) distribution along the 0.35- nautical-mile (n. mi.) sideline, not including extra ground attenuation.

The maximum PNL is 95 PNdB or less, which would be no problem with current regulations, but it 100- 90 - ElllIII S70 C-, I -8 -6 -4 -2 2 4 6 8 DISTANCE BEHIND AIRPLANE ON 0.35-NAUTICAL-MILE SIDELINE, 103 FT Figure II-23. - Thrust reversal noise on 0.

35-nautical-mile sideline for 300 000-lb CTOL air- plane with four NASA "Quiet Engines"; cascade type reversers.

10 PNdB, and it for compliance with FAR-36 minus could become a marginal problem been Similar calculations have FAR-36 minus 20 PNdB.

would be problem for certainly STOL aircraft. For in-flight augmentor-wing-type made for a 100 000-lb gross-weight the PNL's reversal on the ground, 12 and 13) and for wing-slot core-jet reversal (refs.

illustrates the point made in goal at 500 ft. This well above the 95-PNdB design were a problem at short dis- noise is more of that thrust reverser the preceding paragraph tances.

for STOL air- problem, especially reversal noise can be a significant Since thrust of reducing reverser noise noise levels, methods craft and for decreased certification or shields might be of acoustically treated doors studied. Perhaps the use are being experiments with a target on the sideline. Preliminary used to reduce reverser noise possible shielding benefits.

reverser have indicated of Internal Engine Noise Suppression have been achieved in recent years, In spite of the source noise reductions which of today.

in addition to meet the low noise goals acoustic suppressors are required lower noise goals must be achieved so that the anticipated Even more noise suppression must be located To be effective, the noise suppressors of the future can be realized.

limits the noise source and the observer. This the propagation path between the along of the engine. The suppres- to quieting the internal noise sources use of suppressors these noise sources as shown in figure located in the ducts emanating from sors are II-24.

can be studied in the past few years which Two types of noise suppressors have been been the most exten- of aircraft engines. The first, which has used to reduce the noise liner. In this type of liner, the acoustic is the dissipative acoustic sively investigated, Figure II-24. - Engine with suppression.

energy is directed into an absorptive material on the walls where it is dissipated within the wall material. The dissipation occurs due to shearing of the flow in very small pas- sages or through loss of energy of the jets formed by oscillating flow through the small holes in the wall.

The second type of suppressor, the sonic (or choked) inlet, is essentially a reflec- tive type of device.

The simplest explanation of its operation is that if the steady flow within a duct has obtained sonic velocity then a sound wave cannot propagate against this flow. This implies that this principle can be applied only when the sound is propa- gating against the steady flow. In an actual inlet, however, the mechanism is much more complicated than that implied previously. There will be continuous reflections of the sound wave caused by the varying duct diameter and steady flow Mach number.

Radial and transverse velocity gradients also exist which will refract the sound waves away from the axial direction where they can be swept back away from the inlet by the steady flow. Experimental data indicate a steady increase in suppression as the average inlet Mach number approaches one.

A collaborative NASA/General Electric Company parametric study on choked inlets is underway.

The work involves both acoustic and aerodynamic measurements of a family of 19 different inlet configurations which should provide significant inlet quadrant noise suppression. The tests are being accomplished on a 12-in. -diameter fan, and the hardware represents elements of variable geometry cowl and centerbody systems.

Particular attention is being given to measurements of inlet flow profiles in order to make direct correlations with both the internal and external inlet noise fields.

Both types of suppressors have been tested and have demonstrated that very large noise attenuations can be obtained.

Each, however, has its own unique problems which must be solved before they can be considered as truly efficient suppressors. The dis- sipative suppressors at present require splitter rings for the large attenuations that are anticipated in the future. These rings add to the nacelle weight and to the wetted area which increases the total pressure losses.. These losses will be discussed in more de- tail later.

The choked inlet requires small diffusion angles in order to keep the total pressure losses down to an acceptable level. This requires very long inlets, especially for CTOL engines where wide changes in inlet flow occur between takeoff and approach.

There is also a major problem of the variable geometry, actuation devices, and sliding seals which add to the complexity and weight of the system and compound the control problem.

With continued research, the problems will probably be solved so that both types of suppressors can be used in a more efficient manner. The dissipative suppressors could be used today to obtain large noise attenuations if the performance and weight penalties can be accommodated.

One of the problems in interpreting noise suppressor results from experimental data is illustrated in figure II-25. When a fan noise suppressor is designed for large 10 dB SOUND O PRESSURE - LOW SPEED FAN LEVEL JET NOISE- - i I liie I I 100 1000 10000 FREQUENCY, Hz CS-63176 Figure II-25. - Noise spectrum reouiring suppression - low speed fan.

attenuations, the fan noise radiated to the far field can be made to be less than some other noise floor such as the jet noise. The jet noise floor is shown here only as an ex- ample. Other sources may include cowl scrubbing noise, turbine noise, combustion noise, casing radiation, or even mechanical accessory noises. In the case of fan com- ponent tests, the installation noise such as the driving device (motor or turbine) may present a noise floor.

Because of these noise floors, there has probably never been a test of a complete engine with large amounts of acoustic suppression in which the total magnitude of the suppression could be evaluated.

An example of this point can be found in the results from the NASA Lewis quiet engine tests to be discussed later.

The suppressors for this engine were designed to produce about a 15-PNdB noise reduction. It was anticipated in advance that these decreases could never be reflected in corresponding overall engine noise levels, but the suppressors were designed so that this large suppression could possibly be observed on a fan component test.

However, a 7-PNdB noise reduction was observed with the engine test. At takeoff, the jet noise limited the noise reduction mea- sured, and at approach, the turbine noise was limiting. In view of the previous discus- sion, the total suppression effectiveness with regard to a single source cannot usually be determined from tests with a complete engine.

An example of the large attenuations which have been observed with dissipative liners is shown in figure II-26. This is inlet data at 500 from the inlet axis which is less subject to contamination by noise floors than the aft data. Even in this inlet case, however, the suppressed noise data are clearly limited by the jet noise floor up to at least 1000 Hz, and the high speed, quiet engine C data show a turbine noise floor around 8000 Hz.

An example of much larger noise suppression of a high-bypass engine is the TF-34 turbofan engine tests. The acoustic treatment for these engine tests was designed by NASA Lewis and General Electric. The highly suppressed engine is shown in figure 500 ANGLE TAKE-OFF POWER - -o- - A 1C SOLID SYMBOLS, SUPPRESSE 90- SOUND PRESSURE LEVEL 85 AT 200 FT SIDELINE, I I I FREQUENCY, Hz Figure II-26. - Comparison of engine inlet spectra.

FAN EXIT SUPPRESSORS / A,\ CORE FAN INLET /SUPPRESSORS SUPPRESSORS / I /

I /

Figure II-27. - TF34 fully treated nacelle separate flow.

II-27. The design incorporated three treated splitter rings and treated outer duct wall at the fan inlet. The aft fan treatment consisted of two treated rings with both inner and outer walls of the duct also treated. The aft fan duct wall treatment consisted of three different thicknesses to suppress a broad range of frequencies. The core exhaust duct consisted of a bulk absorber material of two different thicknesses.

The objective of the tests was to obtain noise data for a heavily suppressed, high- bypass turbofan suitable for a STOL externally blown flap (EBF) aircraft. The design fan inlet suppression was about 30 dB and the aft fan suppression 37 dB. The design core exhaust suppression was about 25 dB. The results of the tests indicated an overall noise reduction of 23 PNdB with a reduction of 35 dB in the far field for the fan blade passing frequency. Fan pure tone narrow band noise reductions were as high as 50 dB in the far field (fig. II-28). Performance losses due to the suppressed nacelle were about 15 percent. Calculations made for splitter designs that have been optimized for aerodynamic performance indicate that the total takeoff thrust loss at maximum power should be approximately 8 percent.

100 - UNSUPPRESSED SOUND 90 PRESSURE LEVEL dB 70 - 60 - SUPPRESSED

so 111 1

0 2 4 6 8 10 FREQUENCY, kHz cs-6475so Figure II-28. - TF-34 noise spectra at 100 ft. SLS fan speed, 7140 rpm; angle from inlet, 600.

The suppressors for the NASA quiet engine will be tested further at Lewis. An at- tempt will be made to determine the actual noise attenuation provided by these suppres- sors. This will be done by using internal noise measurements and fan-alone tests.

Theoretical work on noise suppressors is continuing to provide a better understand- ing of dissipative suppressors and to provide better design techniques. This will involve both in-house efforts and outside grants and contracts. The theoretical effort will be both in analytical extensions as well as a new numerical technique which has recently been initiated.

These extensions should provide a more realistic model for the sound propagation in suppressors with cross-sectional variations and steady flow gradients.

It is anticipated that by using the velocity gradients in an ingenious manner the suppres- sors can be made much more efficient and thus reduce the aerodynamic losses.

Experimental studies with sonic (or choked) inlets have been conducted at Lewis and - at Boeing-Seattle under NASA contracts. The Lewis work, which has used a 5 -in. fan in the past, will be extended by extensive testing with a 12-in. duct in the near future.

The Boeing studies have used a 12-in. fan and the JT3D engine.

The results of these studies have demonstrated that the choked inlet can provide very high attenuations of the forward radiated fan noise while incurring only small losses, provided the diffusion angles are kept sufficiently small. Similar results were obtained at NASA and Boeing and can be summarized as follows. With a translating centerbody inlet at static ambient conditions, the noise reduction obtained was 29 dB at the blade-passage frequency with large reductions throughout the noise spectrum. These results were accomplished with only a 1.5 percent loss in total pressure. The low- pressure losses quoted for the choked inlet are a result of using low diffusion angles be- tween the inlet throat and the fan face.

If the flow variations between cruise, takeoff, and approach power settings are very small (as in STOL applications), then these diffu- sion angles can be accommodated in a reasonable length of inlet. However, for CTOL aircraft, the flow variation (and thus area ratio variation) is quite large and the flow in short inlets will separate and cause large total pressure losses. The long inlets re- quired will result in increased friction losses along the wetted surfaces and will increase the weight of the nacelle.

With an ambient velocity of 150 ft/sec and an angle of attack of 350, the noise reduc- tion at the blade-passage frequency was 22 dB again with only a 1.5 percent loss in total pressure. However, in addition, the airflow distortion at the face of the fan tends to in- crease with both inlet Mach number and angle of attack. This is a factor which must be given careful consideration with regard to fan stall limits. Both the static and crossflow noise measurements were limited by noise floors at the maximum flow rate. The length- to-diameter ratios of these inlets were about 1.

A full-scale, choked inlet will be tested on quiet engine C at Lewis.

This is a CTOL engine and will require a long inlet. Scale-model tests will be made on a 12-in.

inlet at Langley Research Center in an anechoic chamber to select the L/D for the full-sized in- let.

The L/D variation will be from 1.4 to 1.85 on this scale model.

These models will later be tested at Lewis in a wind tunnel with cross flow.

The theoretical work mentioned earlier in connection with dissipative liners will also be useful in obtaining a better understanding of the acoustic performance of an ac- tual choked inlet with its axial, radial, and transverse velocity gradients.

The large noise attenuations obtained with dissipative suppressors are not obtained without performance penalties.

A continuing effort is being maintained to reduce these losses without giving up the noise attenuation.

In any application, a decision will have to be made on the delicate balance between what goals are desirable and what economic penalties are acceptable.

The perforations necessary in any lining material will in- crease the skin friction over that of a smooth plate. Current estimates are an increased skin friction for perforated sheet metal of 50 percent and for woven structures about 85 percent.

The addition of splitter rings in an engine nacelle increases the losses in three ways: (1) the wetted area of the added acoustic material on the splitter rings in- creases the friction-drag loss; (2) the rings are immersed in high-velocity air and have form drag; and (3) the available flow area is reduced by the cross-sectional area occu- pied by the rings, thereby increasing the steady-flow Mach number. This increases the total pressure losses as the square of the Mach number. The Mach number can be kept down by expanding the nacelle cross section, but this then increases the external drag at cruise. The increased weight of the nacelle also represents an economic loss since it reduces the payload.

1<fHIc <2 10 - 0.6 10 - BOEING NACELLE 0.6 ESTIMATES O CRUISE .4 O TAKEOFF PRESSURE LOSS, 5- AP P % Il l 0 5 10 15 20 SOUND ATTENUATION, APNdB CS-63424 Figure II-29. - Estimated pressure loss due to suppression.

The total pressure loss versus noise reduction can be illustrated as in figure II-29.

As discussed earlier, this illustration may be somewhat pessimistic, but it certainly il- lustrates the trends correctly.

Small initial noise reductions (5 to 10 PNdB) can be ob- tained quite efficiently.

The baseline engine will have blade-passage tones and harmonics and even low frequency multiple-pure tones standing out beyond the broadband noise level.

These tones tend to concentrate near the outer wall and are very susceptible to outer cowl treatment only. The only penalties involved in a moderate reduction of these tones are the increased skin friction and possibly additional weight of the outer cowl. The per- ceived noise level is reduced almost directly with the reduction of dominant tones in the frequency range of greatest ear sensitivity (3000 to 4000 Hz). However, once these dominant tones have been reduced, other parts of the spectrum become dominant and a much broader frequency range must be attacked by the suppressor. This requires the use of several different wall structures to reduce these wide ranges of frequency and ultimately the use of splitter rings if the attenuation required is large. When this oc- curs, the pressure losses begin to rise rapidly as shown in figure II-29. It is apparent that there is considerable incentive to make these dissipative suppressors much more effective than present technology indicates.

Present research efforts are directed at making noise suppressors much more effi- cient. Emphasis is being placed both in theoretical and experimental programs.

The theoretical models are being extended to treat noise propagation with dissimilar wall constructions, area variations, and steady flow velocity gradients. It is anticipated that better use of reflection and refraction phenomena will greatly increase the efficiency of noise suppressors.

The use of axial and radial velocity profiles within a dissipative suppressor leads to a type of hybrid suppressor.

This hybrid utilizes the best properties of both the dissipa- tive liner and the choked inlet. The velocity gradients might be used to refract the sound waves into the wall where they might be absorbed. Perhaps a suppressor without splitter rings, hard choking, or variable geometry may be realized with such a configu- ration. Preliminary data taken by General Electric on a high velocity inlet suppressor indicates the approach has merit. A hybrid suppressor will be tested at NASA Lewis Research Center.

Engine System Noise The preceding discussion has dealt with several components of engine noise. The concern, however, is with the noise that we hear, that is, total engine noise.

What we hear today is generally fan noise.

It is currently the loudest single noise source and thus it dominates in establishing the noise of the total system. Progress has been made in reducing fan noise and further progress will lower the total engine system noise, but only to a point at which another contributor, such as core-jet noise, becomes dominant.

By increasing the engine bypass ratio, we can (within certain cycle limitations) reduce the jet noise by extracting more energy from the core jet, thereby reducing its velocity.

Coupled with other jet-noise-suppression techniques, the overall engine noise level may then be reduced until another source of noise becomes dominant; perhaps turbine noise or possibly fan noise again will become dominant and set the total noise level. It is likely that we are today progressing in our research to the point at which combustor and turbine noises will shortly become the determinants of total engine noise. These are accordingly moving up in priority for future research effort.

Reducing the total system noise of an engine also involves several complex inter- actions. For example, if higher bypass ratio is used to reduce core jet noise, the tur- bines may have to be more highly loaded, thereby increasing the turbine noise. Such interactions must be carefully weighed in the engine design to achieve a best compro- mise and a minimum noise for the total engine.

In a similar vein, the performance or economic penalty of each potential noise reduction feature must be carefully balanced to be gained.

against the noise benefit problems, coupled with recognition of these superimposed effects and interaction In and economic compromises involved, NASA has undertaken the the many judgments the noise reduction features building and testing of complete research engines wherein in a real and total environment. In this way any unexpected are required to function that the industry may incorporate difficulties may be brought to light and solved such such features in future engines with reasonable risk. The NASA quiet engine, refan, are examples of this approach.

and quiet clean STOL experimental engine (QCSEE) NASA Quiet Engine Program initiated about 5 years ago with the objectives The NASA Quiet Engine Program was in engine tests the of developing engine noise reduction technology and demonstrating that this technology would have on reducing engine noise. An additional combined effect economics resulting from the mea- objective was to determine the impact on airplane sures necessary to reduce the noise.

quiet engine program are shown In figure II-30 the major elements of the overall along with a schedule.

The engine design definition studies determined the engine cycle, mechanical arrangements, and other characteristics required to achieve the noise goals initiated in mid-1969 set for the quiet engines. Following these studies a contract was for the design, fabrication, and testing of two quiet engines. In parallel with the engine a contract with the Boeing Company was initiated to provide NASA with an program acoustically treated, flight-type nacelle for testing on one of the quiet engines. The acoustically treated nacelle was tested at NASA Lewis, as part of the in-house program, to determine quiet engine noise levels with full fan noise suppression.

CALENDAR YEAR 1967 1968 1969 1970 1971 1972 1973 1974 ENGINE DESIGN DEFINITION STUDIES (ALLYSON, PIN) (GE): QUIET ENGINE DESIGN 3 FABRICATION TEST ACOUSTIC NACELLE (BOEING): DESIGN FABRICATION QUIET ENGINE TESTING NASA-LEWIS: ENGINE A ENGINE C Figure II-30. - Quiet engine program.

The turbofan engine, which is the type commonly used in the current transport fleet, has two major noise sources. They are jet noise and turbomachinery noise.

The noise reduction features incorporated into the quiet engine designs to counteract these two noise sources are shown by the following: Jet noise: (1) High bypass ratio 5-6, gives low jet velocity and low jet noise Fan source noise: (1) Large spacing between fan rotor and stator, reduces interaction noise (2) Low tip speed (1160 ft/sec), reduces fan noise production (3) High tip speed fan (1550 ft/sec), requires additional suppression for low noise but improves engine weight (4) Optimum ratio of fan stator to rotor blades Fan noise suppression: (1) Sound absorbing liners in fan inlet and discharge ducts A high bypass ratio engine was chosen to reduce jet velocity and consequently jet noise.

A number of features were incorporated to reduce fan noise production.

A relatively large rotor-stator spacing of two rotor chords was employed to reduce fan discrete frequency noise.

A choice of rotor tip speeds was available for the fan design.

Low tip speed fans have been found to produce less noise, while high tip speed fans can improve airplane economics by reducing engine weight, but they require additional noise suppres- sion to achieve equally low noise output.

Both approaches were evaluated in this pro- gram.

Finally, a noise governed optimum ratio of number of fan stator to rotor blades was employed.

The ratio was 2.25.

In addition to design features aimed at low fan noise production, the fan noise can be reduced further by the addition of sound absorbing liners to the inlet and outlet ducts. This was also investigated on the quiet engines.

A cross section of quiet engines A and C with full fan acoustic treatment applied is shown in figure II-31. Also shown are some of the important performance and design characteristics of the engines. Both engines were designed to produce 22 000 lb of thrust, and this puts them in the thrust class of the JT3D engines used in the DC-8 and 707 type aircraft.

Engine C, the high-speed engine, has a single stage fan with a design fan tip speed of 1550 ft/sec, while engine A, the low speed engine, has a single stage fan with a tip speed design point of 1160 ft/sec. To obtain a major cost saving, both engines use the CF-6 engine core, and for this application it is oversized; therefore, the en- gines are not flight weight.

The low speed turbines on the CF-6 cores had to be modified as indicated to meet the power and speed requirements of the fans.

The acoustic performance of the quiet engine was determined for both unsuppressed and suppressed conditions. The results of the baseline, or unsuppressed, engine tests ENGINE C ENGINE A ENGINE C FAN PRESSURE RATIO 1.5 1.6 BYPASS RATIO 6.1 5.1 THRUST, LB 22000 22000 ENGINE CORE CF-6 CF-6 FAN TIP SPEED, FTISEC 1160 1550 Figure II-31. - NASA-GE quiet engines with full suppression.

100 H

PERCEIVED 90 0 ENGINE A NOISE 0 ENGINE C LEVEL, 80 - PNdB zo I I I I I I I 0 20 40 60 80 100 120 140 160 ANGLE FROM INLET, DEG Figure 11-32. - Baseline engine perceived noise directivity.

Approach speed, 370-ft sideline.

CS-63384 100 - PERCEIVED go - NOISE 0 ENGINE A LEVEL, 80 - 3 ENGINE C PNdB 701 I I1 0 20 40 60 80 100 120 140 160 ANGLE FROM INLET, DEG Figure 11-33. - Baseline engine perceived noise directivity.

Takeoff speed, 1000-ft sideline.

CS-63390 are shown in terms of perceived noise in figures II-32 and II-33. Figure II-32 gives data for the approach engine speed while figure II-33 represents takeoff engine speed.

These data show that the noise levels of both engines are quite similar except for the takeoff speed in the front end of engine.

Here the high-speed engine noise is about 7 PNdB higher, and this increase is a result of the "multiple pure-tone" noise asso- ciated with the supersonic operation of the fan blade tip.

Since the baseline engines are dominated by fan noise, adding fan suppression in terms of wall and splitter acoustic treatment (see fig. II-31) significantly reduced engine perceived noise levels.

However, engine C, the high speed engine, requires additional acoustic treatment in both the fan ducts and the core nozzle in order to obtain noise levels as low as engine A. This is a result of the higher fan noise and also a higher tur- bine noise level associated with the high engine speed operation.

Projections of the quiet engine ground static test data to aircraft flyover conditions were made to compare these results with current aircraft. Since the thrust of the quiet engines is in the class of the engines used on the DC-8 and 707, the comparison was made using the DC-8 airplane. The noise results are expressed in terms of effective perceived noise level (EPNdB) which is the noise unit used by the FAA for regulating aircraft noise. The results are shown in table II-1. Measurements of the DC-8 aircraft TABLE ]1-1. - FLYOVER NOISE COMPARISON Takeoff Approach EPNdB DC-8 116 FAR-36 104 106 Baseline quiet engine A 97 98 Quiet engine A with acoustic nacelle 90 89 noise at the takeoff and approach measuring stations show noise levels of 116 and 118 dB, respectively.

These numbers compare with FAR-36 limits of 104 and 106, respectively.

The baseline (unsuppressed) quiet engine A is seen to be some 7 to 8 dB lower than cur- rent FAR-36 regulations and about 20 dB below the DC-8. Furthermore, the addition of an acoustically treated nacelle lowers the noise levels another 7 dB.

These results clearly indicate that the potential for lower noise levels of future aircraft is good; how- ever, it is also necessary to consider the impact that lowering noise levels will have on aircraft economics.

A study of this nature was performed using a 200 000-lb gross weight trijet as a typical aircraft. The results of this study are shown in figure II-34.

Here direct oper- 12 - QE-A TECHNOLOGY --- QE-C TECHNOLOGY 10 - SPLITIERS 3 INLET & 2 EXHAUST A DOC, 6 - 4 - '-1 EXHAUSTSPL z PL-S 7 LrR 2 DUCT WALLS a rUNSUPPRESSED ' "-SPLITTERS . 1 INLET & I EXHAUST I I I 0 . - 5 FAR 36 -5 -10 -15 NOISE LEVEL, A dB Figure II-34. - Comparison of quiet engine A and C economics with variable fan suppression 200,000 Ib, tri-jet ating cost (DOC) is plotted against aircraft noise level relative to FAR-36 noise regula- tions for both the high-speed and low-speed engines. The curves shown for each engine represent various degrees of fan acoustic treatment starting with an unsuppressed case on the lower end of the curves and ending with wall treatment plus three inlet and two exhaust splitters at the upper end. The higher speed engine is more economical (-2. 5 percent DOC) in an unsuppressed condition because the high engine speed allows the number of turbine and compressor stages to be reduced, thereby reducing engine weight.

However, it produces more noise as stated previously. The knee in the curves (where DOC begins to increase rapidly with noise reduction) results from increased engine weight and engine pressure losses that accrue as acoustic splitters are added to the fan inlet and exhaust ducts. As a result, the low speed engine (A), even though it is a basic- ally less efficient engine, is more economical as lower noise levels are reached. The cost of obtaining a noise level of FAR-36 - 10 PNdB, using the A type engine, is seen to be about 4 percent in DOC.

It is obvious, however, that to progress beyond the FAR-36 - 10 PNdB noise levels economically a vigorous noise reduction technology program is required. Advances in noise source reduction and improved suppression efficiency are areas of major impor- tance for future technology programs. The fan and possibly the turbine are the primary candidates for source noise reduction programs. Improvements in suppression tech- nology are needed to increase acoustic treatment effectiveness so that less treatment will be required for a given noise reduction and also to reduce the weight per unit area of treatment by incorporating new materials or fabrication concepts or both.

The use of a sonic inlet also is a promising technique for reducing the cost of noise suppression.

This concept will also be evaluated in future programs.

NASA Refan Program The Refan Program applies current source abatement technology to the engines that power the narrow-body aircraft in the U.S.

civil fleet.

These narrow-body aircraft (707, 727, 737, DC-8, and DC-9) introduced into service beginning in the late 1950's are the noisiest aircraft in the civil fleet. Previously developed noise reduction tech- nology is being applied to the propulsion systems of these aircraft.

No advances in the state-of-the-art are anticipated.

The program objectives are to demonstrate through development of retrofit kits that the noise produced by the narrow-body fleet can be re- duced to 5 to 10 EPNdB below FAR-36 while retaining demonstrated engine reliability and maintainability, causing no degradation of aircraft performance or safety, and all at an acceptable fleet retrofit cost.

Close coordination of the program is being maintained with the Department of Transportation through the Joint DOT/NASA Office of Noise Abatement.

This office is using input data from both this program and from the FAA Acoustic Nacelle Program to assess the costs and benefits of the retrofit options.

Four approaches or combinations of approaches are possible for reducing the noise exposure in the near-airport environment using currently available technology: (1) Retirement of the narrow-body fleet (2) Nacelle acoustic treatment (3) Engine and nacelle modifications (4) A completely new engine Early replacement of the narrow-body fleet with quieter wide-body aircraft to main- tain the present capacity is estimated to cost on the order of $8 billion.

Thus, early retirement of the narrow-body fleet appears to be prohibitively expensive.

Considerable technology effort has been applied to a completely new engine through the NASA Quiet Engine Program.

Development of a completely new quiet engine, however, is being con- sidered primarily for new aircraft.

This option would be extremely costly, particularly in view of the limited life remaining in these aircraft.

A new engine would also prove difficult to install particularly on the JT8D-powered aircraft - 727, 737, and DC-9.

An extensive technology effort has also been applied to nacelle treatment through early NASA work on acoustic nacelles for the 707 and DC-8 aircraft and through the current FAA program in which all five narrow-body aircraft are being considered.

This ap- proach is limited in the amount of noise reduction achievable, particularly during take- off, and it results in some performance degradation of the aircraft.

The NASA Refan Program deals with the third option - engine and nacelle modifica- tions. This approach is attractive since it would not result in the performance loss as- sociated with nacelle treatment only, would be less costly than a completely new engine, and would provide substantial noise relief to the airport community.

It also might afford a convenient opportunity to install a retrofit combustor on JT3D engines to reduce smoke and other emissions while the aircraft are undergoing a refan modification.

The JT8D engines already have a "smokeless" combustor that has been retrofitted to U.S. air- craft.

Technical approach - The objectives of the NASA Refan Program are to be accom- plished by developing retrofit kits which when installed will result in significant reduc- tions of the two main sources of engine noise. These sources are the turbomachinery noise generated by the interaction of the air with the rotating and stationary blade rows, principally in the fan stage, and the jet noise generated downstream of the engine where the exhaust jet mixes with the surrounding atmosphere.

Fan turbomachinery noise is reduced by use of a single-stage fan with greater spacing between the rotating and stationary stages than exists in the current two-stage fans, and by proper selection of the number of rotor blades and stator vanes for mini- mum noise generation.

The jet noise problem is attacked by reducing the velocity of the jet exhaust. This is accomplished by extracting more work from the turbine to drive a larger diameter fan. The larger fan results in a larger bypass-ratio engine which generates slightly more thrust at lower jet exit velocity. Additional noise reduction beyond that obtainable from engine modifications will be obtained by lining the internal surfaces of the engine nacelle with sound absorbing material and possibly by using acoustically treated inlet splitter rings.

History - The Refan Program was initiated in August 1972. The scope of the pro- gram encompassed noise and pollution reduction for the JT3D and JT8D engines. The JT3D engine powers the 707 and DC-8 aircraft and the JT8D powers the 727, 737, and DC-9 aircraft. Phase I contracts were let for design and analysis of the engine and nacelle modifications with three major contractors: Pratt & Whitney Aircraft, a Divi- sion of United Aircraft Corporation; The Boeing Company; and the Douglas Aircraft Company, a Division of McDonnell Douglas Corporation. Small contracts were also let with American Airlines and United Air Lines for consulting work to assure that the modifications being considered incorporate as many of the user airlines' requirements as possible.

In January 1973, program funding curtailment forced limiting the scope of the pro- gram to only one engine. The joint NASA/DOT/FAA decision was to proceed with the JT8D rather than the JT3D. The basic reason for this decision was that the JT8D- powered aircraft will have a larger impact on the aircraft noise exposure in the 1980's.

There was no technical reason for deferring further work on the JT3D. The modified JT3D-9 engine appeared to be a low technical risk development.

The design definition that was completed on the engine and the DC-8 and 707 installations had indicated no sig- nificant problems to be expected in implementing a retrofit of these aircraft.

Program status - Currently the three major contractors are progressing on sched- ule with the design of JT8D modifications and installation designs for use of the refanned engines on the 727, 737, and DC-9 aircraft.

The overall engine design modifications have been selected and are shown in cutaway compared to the existing JT8D engine in figure II-35.

The two-stage fan is replaced with a single-stage fan with approximately a

REFAN

ACOUSTIC

TREATMENT

SINGLE -STAGE FAN

NASA W47 NASA NQ R73-15866 1 30-73 Figure II-35. - JT8D engine.

two-chord spacing between the rotor and exit guide vanes.

The core engine pressure ratio and flow capacity are maintained by inserting two core booster stages in front of the low pressure compressor to compensate for loss of the pressure-producing capabil- ity of the second fan stage in the core region.

Because of the additional swirl of the exit flow caused by the increased work output of the turbine, a slight recambering of the fourth stage turbine rotor blade is required to reduce the swirl angle. The acoustic treatment is shown schematically.

Trade studies are being conducted by the airframe contractors to determine the optimum amount of acoustic treatment to use considering noise, aircraft performance, and cost.

The effects of these engine modifications on some key engine parameters are shown in table II-2 where the refanned JT8D is compared to the current engine.

The larger di- ameter fan results in more airflow, a larger bypass ratio, and an increase in fan tip speed.

The higher tip speed will generate noise in the form of multiple pure tones which previous technology development efforts show can effectively be suppressed with acoustic treatment.

The engine cycle temperature has been selected to remain unchanged in the TABLE 1-2. - JT8D ENGINE CHARACTERISTICS Current JT8D-9 Refan JT8D-109 49.2 in. 40.5 Fan tip diameter, Sea level airflow, lb/sec 319 1.05 2.04 Bypass ratio 1420 1600 Fan tip speed at takeoff, ft/sec 1880, Cycle temperature, oF jet velocity, ft/sec 1766 1479 Primary interest of maintaining the same engine reliability and life. Primary jet velocity has been lowered approximately 300 ft/sec. This velocity reduction is the key to the de- crease in jet mixing noise. Mixing of the core flow more effectively with the fan bypass noise even more.

flow would lower the resultant exit velocity and jet Wind tunnel tests have been completed on DC-9 models and are in progress on the center duct for the 727 aircraft. No installation problems resulting from use of the larger refanned engines have been uncovered in these tests. In fact, wind tunnel model tests on the DC-9 show that the larger nacelles result in a decrease in airplane drag be- cause of a more favorable flow conditions induced on the upper surface of the wing. Deep stall recovery capability, a serious consideration on the DC-9, appears to be adequate after installation of refanned engines and no modifications are anticipated for the tail control surfaces.

Aircraft performance with refanned engines depends greatly on the amount of acous- tic treatment employed. Table II-3 presents a list of parameters for the 727 aircraft and the range of changes that can be expected when the 727 is retrofitted with refanned engines. Not until a final nacelle configuration is selected and more component tests are completed can firm performance numbers be quoted. However, in general, it can TABLE II-3. - 727 AIRCRAFT PERFORMANCE CHARACTERISTICS Current Refan, percent change Installed takeoff thrust per engine, lb 12 700 5.7 to 10.0 Cruise thrust specific fuel consumption, 0.83 -3.4 to +2.4 (lb/hr)/lb Operating empty weight, lb 99 000 2.0 to 4. 1 FAR takeoff field length, ft 8370 -3.5 to -8.4 Range at 30 000 ft (M = 0.84), n. mi. 1355 a0 to -20.

aRecertified for higher gross weight.

be said that with refanned JT8D engines, aircraft takeoff thrust will be increased; spe- cific fuel consumption at cruise will be slightly better or slightly worse than for current aircraft depending on the amount of nacelle acoustic treatment used; the operating empty weight will be greater due to heavier engines and nacelles; takeoff field length require- ments will be less due to the increased thrust; and aircraft range for the vast majority of the current route structure operations will be unaffected.

If range loss should be a problem for certain operators using these aircraft at maximum range, aircraft re- certification at a higher gross weight to recover the range with.additional fuel does not appear to be a problem.

At the time of termination of the JT3D portion of the program, work on the JT3D had progressed into the final engine design stage. The engine layout design has been completed.

The engine and its installation in 707 and DC-8 aircraft are judged to be en- tirely feasible from a technical standpoint.

Engine and airplane characteristics are summarized in tables II-4 and II-5. Presently, the contractors are in the process of writing summary reports on the JT3D part of the program.

Acoustic performance of the aircraft retrofitted with refanned engines has been esti- mated and updated continuously throughout the course of the program. The improve- ments in noise will be discussed in a later section on benefits.

TABLE II-4. - JT3D ENGINE PERFORMANCE CHARACTERISTICS Current JT3D-3B Refanned JT3D-9 Fan tip diameter, in.

50.2 56.6 Sea level airflow, lb/sec Bypass ratio 1.36 2.25 Fan tip speed at takeoff, ft/sec 1423 Cycle temperature, OF 1703 Primary jet velocity, ft/sec 1580 TABLE II-5. - 707 (JT3D) AIRCRAFT PERFORMANCE CHARACTERISTICS Current JT3D-3B Refanned JT3D-9 Performance change, percent Installed takeoff thrust per engine, lb 15 370 16 720 +8.8 Cruise thrust specific fuel consumption, 0.85 0.82 -3.2 (lb/hr)/lb Operating empty weight, lb 145 000 148 920 to 149 560 2.7 to 3.1 Maximum taxi weight, lb 000 336 000 to 341 050 0.0 to 1.5 FAR takeoff field length, ft 11 350 10 250 to 10 680 -9.7 to -5.9 Range, n. mi.

4630 to 4770 -2.9 to 0.0 with plans and schedule - Current Refan Program plans are to proceed Program 737, .and DC-9 aircraft consistent with development of JT8D retrofit kits for use on 727, II is $40 million funding authorization through FY 1975. The program scope for phase a of the decreased funding from the earlier plan of $55 presently being considered in view million. Demonstration ground tests and flight tests on one or more aircraft are being considered.

The first major component test will take place in April at Pratt & Whitney where a full-scale fan will undergo acoustic tests. The effectiveness of various amounts of will be demonstrated as well as the overall design concept of the fan.

acoustic treatment The fan to be used was originally designed for the JT3D engine. However, by the addi- tion of inlet guide vanes and some rework of the fan exit guide vanes, the JT8D fan can be closely simulated.

In November, the JT8D fan will be tested together with the low-pressure compressor tun- spool to determine aerodynamic performance. During this same time frame, wind nel tests on inlets and nozzles as well as scale-model fan acoustic tests will be con- ducted.

January 1974 with Refanned JT8D engine ground tests are scheduled to begin in mating to an acoustic nacelle to take place early in the engine ground test program. En- gine performance and acoustic signatures will be established.

Benefits - Obviously the reason for engaging in a retrofit program such as this is to provide a means of improvement in the near-airport noise environment. The Joint Office of Noise Abatement in a cost/benefit analysis is studying the relative benefits and costs associated with various retrofit options. Aircraft noise data at FAR-36 measuring points and aircraft noise levels as a function of slant range from an overflying aircraft are used as input data for their analysis. Table II-6 shows the FAR-36 measuring point data for the current narrow-body aircraft and for refanned aircraft with maximum acoustic treatment. Also shown are the levels required to meet the current noise regu- lations. The refan numbers are the best estimates based on noise data from past expe- rience gained in noise technology programs. They show that substantial improvements in the noise of aircraft, in fact considerably better than FAR-36 requirements, are pos- sible.

A more meaningful measure of the improvements to be realized by refanning is pre- sented when noise footprints on the ground are considered. The area within a noise foot- print contour represents the area which would be exposed to that given noise level or higher. Footprints should be viewed as qualitative comparisons of noise exposure rather than absolute measures of the area or people affected. Figures II-36 and H-37 are 90 EPNdB footprints for the 727 and 707 aircraft, respectively. The current aircraft and aircraft retrofitted with refanned engines in acoustic nacelles are comnpared. Signif- icant reductions in area are evident. Figure II-38 shows 95 EPNdB footprints for the current and refanned 727 aircraft superimposed on O'Hare Airport. The area subjected TABLE 11-6. - AIRCRAFT NOISE LEVELS Aircraft Measurement Baseline Refanned FAR-36 location E PNdB Sideline 102 92 104. 5 727-200 Takeoff/cutback 102 88 Approach 109.5 96 104.5 104 92 103.5 737-200 Sideline 84 96. 5 Takeoff/cutback 96 (advanced) Approach 108 95 103.5 Sideline 101.5 92 103.5 DC-9-32 Takeoff/cutback 97 84 96 Approach 108 95 103.5 Sideline 105.5 DC-8-61 Takeoff/cutback 117 95 103.5 Approach 98 105.5 Sideline 107.5 94 106.5 707-320B Takeoff/cutback 113 95 Approach 119.5 98 106.5 Figure II-36. - 90-EPNdB contours for 727 aircraft 24 2B 32 7 4 - 4 0 12 16 20 -M.

for 707 aircraft Figure II-37. - 90-EPNdB contours Figure II-38 - Noise contours at 0'Hare for 95-EPNdB 727 aircraft this high noise level is greatly reduced.

to event footprints but is also studying The Joint Office is considering not only single of aircraft and the number of the problem on a larger scale considering the fleet mix at 23 major airports throughout the country in terms of a noise exposure fore- operations in the near future, a clearer cast (NEF). As output from this study becomes available picture of the benefits to be gained from refanning will be more precisely quantified.

considered is the overall cost of a retrofit alter- Costs - One factor which must be native and how the cost will be absorbed. This is being considered in depth in the Joint Office studies. Preliminary estimates of retrofit kit costs for refanning have come from cost about $1.4 million the contractors and indicate that a three-engine retrofit would per aircraft, a two-engine aircraft about $1.0 million, and a four-engine aircraft about $1.8 to $1.9 million.

The total cost of a U. S. fleet retrofit with refanned engines is equivalent to the capi- tal that would be generated by a 1 to 2 percent surcharge on tickets over the next 5 to it should be recalled that there is already 10 years. To put this number in perspective, an 8 percent tax levied on tickets used to build a trust fund for airport development.

Summary - The objective of the NASA Refan Program is to develop propulsion sys- tem retrofit kits for the aircraft in the narrow-body fleet. Installation of these retrofit kits would result in substantial reductions in aircraft noise levels and in noise-impacted areas. Funding constraints have forced NASA to defer further government funded work on the JT3D. The layout design of the modified JT3D is complete and preliminary in- stallation designs have been completed.

With the presently planned government re- sources NASA will concentrate future effort on the JT8D.

No technical problems are evident which would hinder development of either JT8D or JT3D retrofit kits. The Joint DOT/NASA Office of Noise Abatement is assessing the costs and benefits of various retrofit options. The refan option is more effective in terms of noise reduction but also is more costly to implement than a retrofit with nacelle acoustic treatment.

A bibliography of pertinent reports on this subject is included at the end of this chapter.

AIRCRAFT NONPROPULSIVE (AIRFRAME) NOISE Airframe noise is defined as the noise generated by an aircraft in flight from sources other than the engine, auxiliary power units, and machine accessories.

Air- frame noise or aircraft nonpropulsive noise sources, as illustrated in figure II-39, thus include noise generated by airflow over the fuselage, wings, nacelles, flap systems, landing gear struts, wheel wells, etc. The NASA and Navy funded an experimental pro- gram (ref.

14) to measure the airframe noise for five small airplanes in the gross weight range 1500 to 39 000 lb during low-altitude power-off flyby operations.

The FUSELAGE VORTICIES TRAILING WING AND WAKES EMPENNAGE LAYER

Oo onBOUNDARY

000000000 0o, 0c- NOSE GEAR WHEEL WELL AND DOORS U FLAPS, SLATS, ETC.

MAIN GEAR WHEEL WELL AND DOORS Figure II-39. - Schematic diagram illustrating sources of aircraft nonpopulsive noise sources.

90- o CONVAIR 8 DOUGLAS DC-3 o 0 (9 80- * AEROCMDR o 1 / * PRUE-2 GLIDER / o CESSNA 150 0 0 0 70- 70 _B - o o 1 0 SPL, dB 60 - 50 - V4 SA SPL= 10 log V '" +11.4 h 2 CL b 40 I I I I I I I I 60 70 80 90 100 120 140 160 180 200 VELOCITY, KTS Figure II-40. - Summary of power-off aerodynamic noise levels for several aircraft as a function of forward velocity in knots. (See ref. 14, ) O MEASURED APPROACH NOISE 0.'W ESTIMATED AIRFRAME NOISE (REF. 16) -PRESENT FAR-36 106-0727-200 0D 0 10-40

OL1011

OLI0ll

0 727-200ADV

102 O0737-200ADV EPNdB 98- LCARD STUDY 1981 GOAL 994- 0 3 600 80OxlO MAXIMUM GROSS TAKEOFF WEIGHT, LB Figure II-41. - Measured and predicted noise levels for several current aircraft.

power-off aerodynamic noise levels which have been normalized to an altitude of 500 ft are shown plotted as a function of velocity in figure II-40. These aircraft were in rela- tively clean condition because of gears and flaps retraction with the exception of the Cessna 150, which had a fixed gear. An empirical relation for noise predictions based on this experimental information was provided and is shown in the lower part of figure II-40. An extrapolation of these data to larger airplanes suggests that the airframe noise can conceivably be the noise floor for future large airplanes for which substantial engine noise treatment has been used. Further data are required on larger aircraft to evaluate this extrapolation. Some of these will be obtained in tests at the NASA Flight Research Center later this year.

In figure II-41 are presented estimated and measured approach noise levels for several current airplanes compared to those specified in FAR-36.

The circle symbols represent certified landing approach noise levels. It is noted that these aircraft gener- ate noise levels which are less than those specified in FAR-36 (from 2 to 5 EPNdB).

The lower dashed curve which is situated 10 EPNdB below the levels of the present noise certification ruling represents the CARD 1981 goal (ref. 15). The hatched region is de- fined in reference 16 as including the airframe noise levels for the aircraft of the figure as estimated by the method of reference 14.

Recent unpublished data from the Lockheed- Georgia Company and the USAF for the C5-A tends to confirm the validity of the hatched region.

The main conclusion of the figure is that the noise associated with all of the nonpropulsive sources (no engine) appear to lie about 10 dB below the noise levels of the present noise certification ruling and could in fact be the limiting noise floor relative to reducing noise levels to and below the 1981 CARD goal (ref. 17), until the nonpropulsive noise is itself reduced.

A program to understand and reduce the nonpropulsive noise is underway at NASA.

Basic studies are planned on full-scale aircraft and complete models and model compo- nents for quiet wind tunnel tests. These studies will provide information relative to the identification and location of airframe noise sources; the manner in which noise varies as a function of angle of attack, local air velocity, turbulence levels, separated flows, etc.; improved prediction methods; and approaches to noise alleviation.

Initial studies will be carried on in quiet wind tunnel facilities such as those in the Langley Aircraft Noise Reduction Laboratory and these studies will eventually identify concepts for full- scale testing on aircraft and gliders.

The components of airframe noise are judged to be associated with turbulent flows and flow separation and hence prediction methods will inherently involve empiricism.

Empirical constants used in the predictions will be evaluated in experimental studies.

At this time the problem of scaling frequencies and amplitudes of disturbances associ- ated with unsteady flows is not well enough understood.

It is believed that a series of measurements of the fluctuating surface pressures on airfoils and fuselage under a vari- ety of inflow conditions and for known localized flow situations will be required. To de- needed for prediction, some tests in con- velop systematically the empirical constants the prediction of ventional wind tunnels at varying model scales may also be required for pressure patterns. Measurements are required in such studies to the fluctuating surface as well as the surface flow patterns and the fluctuating flow patterns, define the mean pressure fluctuations.

planned in quiet wind tunnel facilities as well as conven- Selected experiments are quiet wind tunnel facility tional facilities to establish validity of the data. In addition, tests will be planned to correlate the surface pressure fluctuations with the radiated flow situations.

sound field for a variety of fuselage and wing flap AIRCRAFT SUPERSONIC TRANSPORT The supersonic transport (ST) has two associated environmental noise problems that are of serious concern with regard to public acceptance. These are the noise from the during landing, takeoff, and ground operations at airports and the sonic booms engines resulting from flight operations over inhabited areas at Mach numbers higher than 1.0.

These problems are different in nature and require different approaches for noise con- trol.

Contained herein are discussions of jet engine noise control approaches involving variable engine cycles and such suppression devices as multielement nozzles, ejectors, reflectors, and absorbers. The current state of knowledge regarding sonic boom con- trol by configuration shaping and operational procedures is also summarized. Planned research in both areas of work is indicated where appropriate.

JET NOISE ABATEMENT TECHNOLOGY A variety of engine cycles can be considered for ST aircraft. However, the jet ex- haust velocities for the cycles are considerably higher than those for CTOL aircraft. As a consequence, the jet noise for these engines, being a primary function of jet velocity, is much louder than for those used in CTOL aircraft. Although the fan for ST aircraft engines generally operates at low bypass ratios and high pressure ratios with resultant high noise, the unsuppressed jet due to its high velocity is the dominant noise source.

There is, therefore, a need to suppress jet noise in order to render supersonic trans- port aircraft acceptable to the community. As the jet noise is suppressed the fan noise and core (internal) noise may become dominant, as in the case for CTOL aircraft. Noise attenuation means for these noise sources with ST aircraft are similar to those applied to CTOL aircraft.

Variable Engine Cycles The use of variable cycle engines has been proposed in order to help reduce jet noise.

Some conceptual examples of these engines are shown in figure II-42.

In figure II-42(a) a low-noise auxiliary fan augments the thrust of the main propulsion system, a duct-burning turbofan, at takeoff. Fan air from the main engine powers the auxiliary fan. The power extracted from the turbofan results in a low jet noise for the main en- gine also. After takeoff the auxiliary fan is not operated.

In the augmentor wing concept (see also section POWERED LIFT AIRCRAFT, p. 86) (fig. II-42(b)) bypass air from a turbofan is ducted through the wings at takeoff and ejected at high velocity into a two-dimensional (split-flap) ejector. Ambient air is en- trained by the pumping action of the ejector and mixed with the fan bypass air. This re- sults in a much lower jet velocity at the ejector exit plane and consequently the mixed flow noise level is much reduced.

The mixing noise of the nozzle flow with the entrained air can be reduced significantly by suitable nozzle design (multielement nozzle) and by acoustically treating the inner surfaces of the ejector. In addition, some thrust augmen- tation due to the ejector also may be obtained.

In the cruise mode the flaps are retract- ed. After takeoff, valves close off the air to the wing ducts and the turbofan is operated (a) AUXILIARY FAN.

(b) AUGMENTOR WING.

(c) FOLDING FAN.

(d) COMPOUND.

(e) SPLIT FANS AND COMPRESSORS.

Figure II-42.

- Variable cycle engine concepts.

in a conventional manner.

turbofan at The folding-fan cycle of figure II-42(c) operates as a high-bypass-ratio the fan blades are removed from the airstream by folding back and takeoff after which the engine operates as a conventional turbojet. The compound engine in figure II-42(d) comprises a central turbofan with satellite turbojets around the fan gas generator. At the thrust takeoff auxiliary inlets supply air to the auxiliary turbojets thus supplementing of the central turbofan. After takeoff the auxiliary inlets are closed and the turbojets fan air. In this mode the engine performs similar to a conventional use supercharged turbojet.

concept in figure II-42(e) uses auxiliary inlets and The split fan and compressor nozzles for the larger airflow at takeoff. As shown in this figure a separate fan is placed in front of and mechanically driven by a turbofan engine. At takeoff forward fan air is exhausted through auxiliary nozzles and the turbofan receives air from the auxil- iary inlets. Thus a much larger airflow is taken on board with a smaller diameter en- gine. After takeoff the auxiliary inlets and nozzles are closed and the air from the for- ward fan goes through the turbofan and the engine operation is similar to a conventional turbofan engine.

From the description of these cycles it can be seen that these engines may be com- plicated and expensive. NASA has contracted with the General Electric Company and Pratt & Whitney to perform analyses of propulsion systems suitable for supersonic transport aircraft. A major goal of the work is to examine systems that can meet severe noise constraints, not only those of today but the possibly more stringent ones of the future. The engine contracts are coordinated with more general studies of the com- plete airplane design being performed by the Boeing, Lockheed, and Douglas airplane companies under contract to Langley Research Center. The Langley contracts will study the technology problems and design tradeoffs for the integrated airframe/engine combination, including such operational constraints as engine noise limits.

Suppression Devices At the high jet exhaust velocities of the more conventional engine cycles considered for a ST aircraft-type engine, the ability to meet acceptable community noise standards is difficult even with the best available jet noise suppressors. Suppression devices that approach present FAR-36 CTOL requirements have been under research; however, the associated aerodynamic performance losses and weight penalties severely compromise the economics of the propulsion and aircraft system. To achieve an environmentally and economically acceptable high-speed transport aircraft will require substantial im- provement in jet noise suppression technology.

Although the approach used in developing jet noise suppressors has been largely empirical, a considerable volume of data has been amassed which is useful in providing guidance in planning new research and in stimulating the development of more general suppression concepts (refs. 18 to 23). The complex nozzle geometries used in noise suppressors (multitubes, chutes, spokes, etc.) lead to complex noise generation and/or reduction mechanisms.

The reduction of high velocity jet noise by means of current suppressor nozzle con- cepts requires the application of four basic techniques, which must be accomplished within a number of constraints. These constraints can be categorized as airplane per- formance (thrust, weight, drag, etc.) and design constraints (mechanical feasibility, installation, retractability, etc.) of the nozzle system. The four noise reduction tech- niques are as follows: (1) Source energy modification. This requires using flow breakup and mixing in such a manner as to establish the desired jet-wake flow patterns to control the noise frequency spectrum, noise source location, and generated noise levels. Multielement nozzles are frequently used for this purpose. Some nozzles of this type tested by NASA are shown in figure II-43.

In this approach the associated thrust, drag, and weight pen- alties must be considered.

(2) Redirection and control of energy generated. This can be accomplished by es- tablishing favorable thermal profiles, additional fluid injection, and mechanical element arrangement.

The objective is to direct the noise at low angles from the jet axis in or- der to maximize the length of the noise transmission path, thus taking advantage of at- mospheric attenuation (item (4)). If most of the noise could be directed in this manner, ground noise levels would be greatly reduced.

(3) Energy absorption. Auxiliary devices such as acoustically lined ejectors, gas sheaths, or reflecting devices (wings, etc.) are used to absorb or intercept the noise once it has been radiated by the jet and prevent the noise from reaching the community.

Redirection and control techniques are used to turn the sound toward the absorption de- vice to take advantage of available attenuation.

(4) Atmospheric attenuation. The transmission of noise through the atmosphere is known to cause attenuation, particularly at the higher frequencies. The noise spectral characteristics can be shaped to take advantage of this effect.

Each of these features must be accomplished while sustaining minimal thrust loss in both the takeoff and cruise modes. The suppression system will generally have to be retracted from the exhaust stream during cruise in order to meet performance and eco- nomic goals. In such cases, mechanical and thermal considerations with regard to the component structure become very important.

The use of multielement nozzles (tubes, chutes, spokes, etc.)

for jet noise sup- pression dates back into the 1950 era. By use of such nozzles the total jet is divided into many smaller jets. Research by industry and government has led to an improved Figure 11-43. - Typical noise suppression nozzles.

Figure II-44 - Experimental high velocity jet noise suppressor nozzle with ejector.

understanding of these devices.

Such devices, however, are accompanied by thrust losses that vary with the particular configuration used. Severe material problems are encountered in placing a suppressor in a hot exhaust stream typical of afterburner oper- ation.

A common type of experimental high-velocity jet noise suppressor is shown in fig- ure II-44.

It consists of a multitube or mixer nozzle combined with an ejector. Such a nozzle system can achieve significant noise reduction and is also potentially able to meet the operational constraints placed upon it.

In this system, the multitubes are used to break up the exhaust flow into many, small jets. This shifts the noise-generation pro- cess to higher frequencies (fig.

II-45) and somewhat lower levels. The reduced noise level at the lower frequencies are associated with the coalesced flow (mixed flow far downstream from the multitube exhaust plane) while the high frequency noise is associ- ated with the small multitubes and occurs near the nozzle exhaust exit plane.

If an acoustically lined ejector shroud surrounds the multitube noise generation region and is designed to absorb noise at the frequencies in this region, it is theoretically possible to reduce the noise to that generated by the lower velocity exhaust jet emerging from the ejector exit plane. This jet noise level is the minimum obtainable with the particular nozzle-ejector system. It should be noted that little of the low frequency noise is atten- uated by the lined ejector since most of it is generated downstream of the ejector exit plane.

The ejector serves as a surface to which an acoustic liner can be attached as well as providing thrust augmentation during takeoff. In order to minimize performance losses in cruise, the ejector can be retracted and the mixer nozzle elements stored within the confines of the engine nacelle. Development of a variable-geometry ejector to minimize thrust losses for the various flight modes can also be considered.

A model and full-scale experimental program has been conducted by The Boeing Company for NASA to acquire a parametric set of acoustically lined ejector jet noise suppression data from which a design technology could be formulated. Some of the noz- zle configurations tested are shown in figure II-46. Ground static tests were conducted using a J-75 turbojet engine with a 37-tube suppressor nozzle.

Ejectors of two lengths, 1 and 2 shroud diameters, were tested with various linings.

A total of 7 lined ejector configurations were evaluated at full scale over the engine operation range of 1.4 to 2.4 engine pressure ratio, corresponding to jet velocities of 1000 to 2000 ft/sec.

Corre- sponding 1/4-scale models of 17 lined ejector configurations were tested over a range of nozzle pressure ratios of 1.4 to 4.0 and jet temperatures from ambient to 15000 F. For both full-scale and 1/4-scale tests a round convergent nozzle was tested to establish baseline acoustic data. The liners were designed using an existing procedure for fan- duct liners.

For a given type of geometry, specified flow Mach number, and desired tuning frequency, corresponding to the frequency of the most annoying noise, the sup- pression was predicted as a function of lining-panel porosity, an example of which is MIXER-TYPE SUPPRESSOR NOZZLE WITH EJECTOR r-STANDARD TAILPIPE SJET NOISE ' s rSUPPRESSOR NOZZLE / JET NOISE SOUND / PRESSURE LEVEL, dB 20dB SUPPRESSOR AND / SHROUD JET NOISE FREO UENCY, Hz Figure II-45. - Jet noise suppression.

00000 000000 0000000 S 0 0 0 0 (a) CIRCULAR NOZZLE. (b) 37-TUBE SUPPRESSOR NOZZLE.

0 0000 0 0 0 0 0 0 0 0 00 000000 0 0 0 0 0 0 00000 (c) 37-TUBE NOZZLE WITH EJECTOR.

Figure II-46. - NASA-Boeing test configurations.

shown in figure II-47.

From a family of such design curves, various linings were chosen to give maximum noise attenuation.

A brief discussion of some of the major re- sults of this study follows.

Figure II-48 shows typical sound spectra for the full-scale tests, at 200 ft from the nozzle and at an angle of 1400 from the nozzle inlet, for various nozzle and ejector con- figurations having 10.6 percent open area panels with a 2. 1-in.

cell depth.

The engine pressure ratio is 2.4, and the jet total temperature 10000 F.

The upper curve is for the round convergent nozzle, showing the characteristic loud, low-frequency noise.

The 37-tube suppressor nozzle shifts the peak noise to higher frequencies and produces some sound pressure level (SPL) reduction at all frequencies.

Adding the 52-in. hard wall ejector gave little further noise reduction; in fact, refraction effects cause higher noise FLOW MACH NO., 0.5 10- TUNING FREQUENCY, 2000 Hz ONE-OCTAVE BANDWIDTH POINT 5 DESIGN ATTENUATION, AdB I I I I I 5 10 15 20 PERCENT OPEN AREA Figure II-47. - Predicted acoustic liner performance.

LINERS: 10.6% OPEN AREA 2.1 IN.CELL DEPTH 120- 110- SOUND NOZZLE CONFIGURATION PRESSURE RU LEVEL AT 100 i-ROUND CONVERGENT 200 FT, ' ,-37-TUBE SUPPRESSOR dB\ 90 - ',-SUPP.

+52-IN. HARDWELL H.W. EJEC.

- SUPP. +52-IN.

LINED EJECTOR '-SUPPRESSOR + 104-IN. LINED EJECTOR 80.

50 100 200 500 1K 2K 5K 10K 113-OCTAVE BAND CENTER FREQUENCY, Hz Figure II-48. - J-75 sound pressure level spectra. Angle measured from inlet, 1400; engine pressure ratio, 2.4.

at some angles for the hard-wall ejector as compared to 37-tube nozzle alone. Adding the lining to the 52-in. ejector yielded a significant reduction in high-frequency noise, with essentially no attenuation at frequencies less than 500 Hz; however, a significant peak remained at a frequency of 2500 Hz, which is in a very annoying range. Doubling the length of the lined ejector provided still further significant reductions in high- frequency noise, with some low-frequency noise reduction also.

These reductions in SPL also yield reductions in the perceived noise level (PNL) as illustrated in figure II-49 for these same configurations.

The reduction of maximum NOZZLE CONFIGURATION 16- 37-TUBE NOZZLE + 104-IN.

NOZZLE + 104-IN. LINED EJECTOR 12 - E8T NOZZLE + 52-IN. LINED PERCEIVED 37-TUBE NOZZLE ALONE NOISE LEVEL SUPPRESSION, A PNdB 37-TUBE NOZZLE+ 52-IN.

HARDWALL EJECTOR 0 111 1000 1200 1400 1600 1800 2000 JET EXHAUST VELOCITY, FTISEC Figure II-49. - J-75 engine PNL suppression at 0.35 nautical mile sideline distance.

PNL below the maximum PNL for the circular nozzle is plotted against nozzle jet veloc- ity. The attenuation generally increases with increasing velocity; this is typical of most suppression devices. The trends are similar to those shown in figure II-48 except for the hard-wall ejector, which gives less suppression than the 37-tube nozzle alone. This is due at least in part to refraction effects, which cause the sound to be directed away from the jet axis, thus reducing the path length over which atmospheric attenuation and spherical radiation can work to reduce the SPL. A maximum suppression of about 15 PNdB was obtained with the 104-in. lined ejector, while the 52-in. lined ejector gave a maximum suppression of about 11 PNdB.

Figure II-50 indicates that 1/4-scale and full-scale 37-tube nozzle data show good agreement in terms of SPL suppression as a function of frequency. The scale model data are shifted in frequency by a factor of 4, in agreement with theory. This indicates the validity of small-scale testing in evaluating suppressor nozzles.

A comparison of the predicted and measured noise attenuation obtained with the 52- in. acoustically lined ejector is shown in figure II-51 in which the sound power reduction is plotted as a function of frequency. While the prediction agrees relatively well with the MODEL SCALE FULL SCALE 15 1 Hz/ 125 KHz 315 'C \ SOUND POWER LEVEL / SUPPRESSION, dB / 0 ............

---- -- --------- -5 100 200 1K 2K 5K 10K 20K 40K 1/3-OCTAVE BAND CENTER FREQUENCY, Hz Figure II-50. - Sound power suppression obtained with 37-tube suppressor nozzles.

Pressure 0 ratio, 2. 4; jet total temperature, F.

MEASURED R CALCULATED 12 12 SOUND POWER REDUCTION, dB /

/ I

4- / /

0I I I \I

I 500 1K 2K 5K 10K 1/ 3-OCTAVE BAND CENTER FREQUENCY, Hz Figure II-51. - Comparison of measured and calculated noise attenuation for J-76 engine with 52-inch acoustically lined ejector.

Pressure ratio, 2. 4 measured values, further improvements in the prediction procedure are needed to ac- count for flow gradients and noise source locations.

The following conclusions were drawn from this liner ejector-suppressor technology study: (1) Lined ejectors provided significant additional jet noise suppression above that provided by a multitube suppressor nozzle by as much as 6 dB.

(2) A maximum sideline perceived noise level reduction of 15 PNdB relative to the round nozzle was obtained with the 2-diameter long acoustically lined ejector.

(3) Further improvements in the acoustic-lining prediction procedure are needed to account for flow gradients and noise source locations.

Under static conditions the 37-tube suppressor nozzle has a constant 2-percent loss in thrust compared with the baseline conical nozzle over the entire pressure ratio range of the engine.

With the ejector, the static thrust of the system was increased at least 8 percent over that of the conical nozzle. This increase in thrust was due in part to the large bellmouth used on the ejector; however, on the basis of other NASA tests with flight-type ejectors (unpublished data), this level of static thrust augmentation also ap- pears reasonable with flight-type hardware.

NASA is also conducting research on novel jet noise suppressors. Shown in figure II-52 is a nozzle design that shows some promise for noise reduction at supersonic ex- haust conditions.

The nozzle is a convergent-divergent nozzle that operates in an over- expanded condition (ref. 24). The design of the nozzle is such that the jet is divided into lobes. An ejector is placed around the nozzle. The convergent portion is a standard eight plates separated by V-gutters. A conical nozzle. The divergent portion consist of Figure II-52. - Divergent lobed suppressor nozzle on J-85 afterburner test rig.

step area increase exists at the nozzle throat.

A low pressure exists in the base cavity formed by this area step.

This low pressure causes the flow to overexpand.

The flow attaches to the plates and is divided among the plates by the V-gutters.

A strong shock structure results, and the jet velocity rapidly decreases.

Some of these effects can be seen in figure II-53, which shows the axial Mach num- ber distribution along the jet centerline and along the plates.

The Mach number is higher along the plate region, which indicates that the flow has expanded into this region and, in fact, is even concentrated in this region.

Shocks exist just upstream of the end of the 2.4- o CENTERLINE 2.0 0 PLATE REGION 1. 6 ..

1. 2 MACH NUMBER NO.

ENDOF .4 - V GUTTER ,-END OF I NOZZLE AXIAL DISTANCE FROM THROAT IN DIAMETERS, x/D Figure II-53. - Mach number distribution along primary plate and nozzle centerline.

V-gutters.

Downstream of these shocks, the Mach number decreases and the flow be- comes subsonic in about 3 or 4 diameters, and the normally long supersonic mixing re- gion has been eliminated.

The particular nozzle tested was optimized for supersonic exhaust conditions, and no noise reduction was obtained at subsonic velocities.

At a typical cold flow operating point for the nozzle with a hard-wall ejector and at a pressure ratio of 3.5, a 14-dB noise reduction in peak noise level and no static thrust loss compared with a conical nozzle was obtained.

Use of a lined ejector should yield a further 5 to 6 dB reduction in noise.

A summary of the current state of the art in supersonic jet noise suppression through the use of multielement nozzles is shown in figure II-54, in which the amount of noise suppression is plotted as a function of static thrust loss.

The plot is based on a pressure ratio of 3.0 across the nozzle exhaust plane and a jet total temperature of 10000 to 20000 F. A reference line is also shown that illustrates the small amount of noise suppression achievable by throttling back an engine in order to operate at thrust.

The shaded regions identified in figure II-54 represent a large mass of data and in- dicate that significant strides have been made in achieving suppression nozzles that have high noise suppression and low thrust losses.

Recent static tests with an advanced con- A PNdB AF 2.0 1.0 ADVANCED RELIMINARY CONCEPTS ESEARCH CURRENT TECHNOLOGY PERCEIVED NOISE REDUCTION, A PNdB "THROTTLE EFFECT I I I 0 8 16 24 THRUST LOSS, AF, PERCENT Figure 11-54. - Current suppressor noise reduction effect- iveness. Pressure ratio, 3; jet total temperature range, 10000 to 20000 F.

5 - BYPASS-TO-CORE AREA RATIO, 1.9 SOUND POWER LEVEL ATTENUATION, AdB -5 - -10 0 .5 1.0 VELOCITY RATIO BYPASS-TO-CORE bypass flow (coaxial - Jet noise attenuation with Figure II-55.

flow nozzle).

cept suppressor utilizing a 61-tube stowable nozzle with an acoustically lined ejector and shield showed an overall sideline noise suppression capability at 0.35 n. mi. of 20 dB with a thrust loss of 5 percent at a jet velocity of 2500 ft/sec (ref. 25).

This suppressor utilized integrated disciplines of mechanical systems, propulsion, and noise.

Since some candidate cycles for a ST propulsion system would produce two exhaust streams, the noise generation in coaxial (bypass-type) nozzles becomes an important problem area. Use of subsonic coaxial nozzles has resulted in noise reduction as shown in figure II-55 (ref.

26). With supersonic exhaust jets (cold flow), still further noise reductions have been reported due to changes in shock-wave patterns (ref. 27). Refrac- tion effects due to temperature differences between the streams may significantly alter the noise directivity.

This phenomenon is similar to that observed when the outer flow RAYS BENT - ROHER ACOUSTIC BY JET AXIS - - TOWARD 0 LE - . _ TEMPERATURE GRADIENT RESULTING REFRACTION INCREASING /-WITHOUT DIRECTIVIlY OF NOISE SPL ~'- WITH REFRACTION 0° 1800 - JET FLOW Figure II-56.

- Refraction of jet noise by hot bypass flow.

stream has a higher speed of sound than the core flow (i. e., helium shield).

It has been suggested that a relatively hot low velocity annular flow as would be generated by a duct burning engine might bring about sideline and flyover noise reduction by bending the sound rays in toward the jet axis as illustrated in figure II-56.

To investigate this phe- nomenon Lewis is modifying an existing cold-flow, coaxial nozzle facility so that either or both streams can be heated to 16000 F at nozzle pressure ratios up to 3 (jet velocities up to about 2500 ft/sec). In addition, it is planned to study contractually noise suppres- sion means for the bypass flow nozzles.

This work will consist of an experimental pro- gram in which the pertinent flow and geometry variables will be studied parametrically.

In-Flight Effects on Suppression Devices So far this discussion has centered on static test results obtained with various sup- pressor nozzles, both aerodynamically (thrust) and acoustically.

The question arises as to how the aircraft forward speed influences the noise suppression characteristics of the various nozzle configurations.

While the understanding of noise generating mechanisms and static suppression techniques have been advanced by research, little effort has yet been expended to develop suppressed jet technology for the final flight application.

Test programs are needed to develop data to aid in understanding flight velocity effects and to be more cognizant of their implications in flight noise predictions.

This becomes particularly important based on the subjective response rating of effective perceived since noise regulations are noise level (EPNL), which is a function not only of peak noise but also its tone content and time duration. Thus, an understanding of flight velocity effects on the directivity of unsuppressed and suppressed noise as well as on the peak noise is required.

NASA and General Electric, under contract to NASA, have performed flight noise unsuppressed nozzles installed on a modified F-106B tests of several suppressed and aircraft. The aircraft is equipped with two auxiliary J85-13 engines, one mounted under each wing (fig. 11-57). Only one auxiliary engine at a time is used for the noise tests.

with F-106B aircraft and 3-85 engines.

Figure II-57. - NASA flight jet noise research tested were shown in figure 1I-43. General trends of flight velocity Some of the nozzles influence applicable to all nozzles have not been observed, indicating the need for further study (refs. 28 and 29).

NASA is conduct- In order to improve the understanding of forward velocity effects, evaluation of the effect of forward velocity on the noise characteris- ing a comparative tics of supersonic-aircraft-type suppressor nozzles both in flight (Lewis Research Cen- in a 40- by 80-foot wind tunnel (Ames Research Center). A calibrated J85-13 ter) and will be installed in the wind tunnel for tests with various nozzles that have already engine flight tested by Lewis. These nozzles will include a 104-elliptical-tube suppressor been acoustically lined shroud. Installation of the en- nozzle (fig. II-58) with and without an gine pod, including the nacelle and inlet, will be as nearly identical as practical for the Mach two programs. Some differences will necessarily exist; for example, the flight 0.33 to 0.40, while the wind tunnel Mach number range number range is approximately is from 0 to 0.30 or less, depending on model drag. Furthermore, for the flight tests, the wind tunnel tests, it the nozzle is in motion with respect to the microphones, and for for jet velocities is not. The comparative evaluation of flyover noise levels will be made C-72-3990 Figure 11-58.

- 104-tube suppressor nozzle with ejector used for in-flight F-106B noise program.

ranging approximately from 1100 to 2100 ft/sec.

The capability of making wind tunnel and flight noise measurements will be further developed and the limitations of each method will be documented.

Using these two ex- perimental techniques, a comparative evaluation of the effects of forward velocity on the noise characteristics of supersonic-transport-type suppressor nozzles will be made.

If the experiments prove sufficiently accurate, the differences in source motion may allow some investigation of source motion effects on noise spectra.

Core Noise The problem of CTOL and STOL core noise was discussed previously.

It was shown that at low jet velocities associated with CTOL core (internal) noise attributed to com- bustion, turbine and flow wakes from structural protrusions (struts, etc.

) emerged as the dominant noise source.

At supersonic jet velocities, however, core noise is a sec- ondary factor when compared to the unsuppressed jet noise.

However, when suppressor nozzles are used, core noise emerges as a floor to jet noise reduction.

A simple sketch is shown in figure II-59 in which pure jet noise and core (internal) noise are related to jet velocities.

At the lower velocities, pure jet noise obeys an 8- power relationship with jet exhaust velocity while at high jet velocities, a 3-power rela- tionship is observed.

Core noise appears to follow a dipole or 6-power relationship with HIGH VELOCITY PURE JET (3RD POWER LAW) - PERCEIVED NOISE LEVEL, '-CORE (INTERNAL) NOISE -LOW VELOCITY PURE JET (8TH POWER LAW) JET VELOCITY Figure 11-59. - Relation of jet and core noise.

jet noise in jet velocity in the region of interest. When a suppressor nozzle is used, the the high-velocity regime is reduced until the core noise level or floor is reached. Core noise is not significantly changed by the use of such nozzles. Reduction of the core noise floor could be achieved by the use of quiet combustors and absorbing some of the remain- of the internal noise floor is difficult ing internal noise by acoustic liners. A reduction to achieve solely with liners because the core noise, particularly combustion noise, is noise. This latter noise is difficult to remove with reason- dominated by low-frequency ably sized lightweight acoustic liners.

SONIC BOOM The sonic boom, which is associated with the shock wave patterns of an aircraft, is of concern only during the phases of the flight which are accomplished at speeds exceed- ing the local speed of sound. The nature of the sonic boom ground exposure patterns are shown in figure II-60. Shown schematically in the figure is an airplane flight track ex- tending from subsonic to supersonic speeds. Beneath the flight track are shown sketches of the shock-wave impingement patterns and the associated distributions of N-wave pres- sures, both along the track and perpendicular to it. In references 30 and 31, informa- tion is presented regarding the state of knowledge of the sonic-boom phenomena dealing with the ontrack and lateral ground exposures during steady flight for quiescent atmo- sphere conditions, the effects of atmospheric dynamics on bringing about sonic-boom- signature variations, and the pressure buildups resulting from accelerated flight.

The significant factors which affect the sonic boom include airplane design, which involves weight, size, and volume and lift distributions; airplane operations, which in- volve altitude, Mach number, and flight path; and the atmosphere, which involves pres- A9 Figure 11-60.- Sonic-boom ground-pressure patterns.

HYPERSONIC SUPERSONIC LOW SUPERSONIC Figure II-61.- Schematic of shock waves associated with three operating regimes.

and turbulence. Discussions of the effects of sure, temperature, and wind gradients they relate to transonic, supersonic, and hypersonic operating each of these factors as regimes shown in figure II-61 are presented.

For Steady Flight and a Steady Atmosphere Sonic-Boom Exposures with the in figure II-62 are the ontrack sonic-boom overpressures (associated Shown in the sketch) as a function of airplane altitude for a num- N-wave shape signature shown transport, and a hypersonic ber of airplanes including a fighter, a bomber, a supersonic transport. The experimental points obtained from overflights of the Lockheed F-104, airplanes (ref. 32) represent averages of a Convair B-58, and North American XB-70 and esti- large number of measurements. The shaded areas represent the predicted on current theory, a standard atmosphere with no wind be- mated nominal values based II-62 indicate that the overpressure decreases ing assumed. The results given in figure with increasing altitude and with decreasing airplane size. Good agreement exists be- conditions.

tween theory and experiment for a wide range of airplanes and operating levels for the hypersonic airplane are justified on the basis of recent The estimated measurements on Apollo 15, 16, and 17 launch and reentry vehicles. Results indicate that satisfactory agreement exists between calculations and the measured boom data

-I-

AP AP 0 FLIGHT DATA N/rn 2 THEORY 200 -lIb/ft

4 e-/////

ESTIMATE - -

3 =4

100 2 50 1 . '4 0 : 0- I I I 13 20 40 60 80 100 120 x 10 ft I I I I 1 1 6 12 18 24 30 36 km ALTITUDE Figure 11-62.- Sonic-boom overpressure characteristics of various supersonic airplanes as a function of altitude.

from the spacecraft at Mach numbers in excess of 10 and at altitudes in excess of 70 kilometers (km). Thus, existing methods involving empirical inputs from wind tun- nels may be useful for predicting sonic boom exposures for very high Mach number and altitude conditions of future hypersonic transports (see ref.

33).

The data in figure II-63 relate to lateral-spread measurements from the XB-70 air- plane at two altitudes and two Mach numbers (see ref.

32). The sonic-boom overpres- sures are shown as functions of lateral distance to either side of the airplane ground track.

The data symbols represent the average of measurements from a large number 2 2 N/m Ib/ft 200 - 4 - THEORY M - 1.5 AT 11.3 km (37 000 ft) 100 - 2 - 4 FLIGHTS .CALC. CUTOFF - , 0 EXPERIMENT 200- 4 - M - 2.0 AT 18.4 km ( 60000 ft) 100 • 13 FLIGHTS 12400 - 40 32 24 16 8 0 8 16 24 32 40 LATERAL DISTANCE FROM GROUND TRACK, n. mi.

Figure 11-63.- Sonic-boom overpressures for the XB-70 airplane as a function of lateral distances for two different flight conditions.

of microphones for 4 flights and 13 flights of the XB-70 at 11.3 km (37 000 ft) and 18.4 km (60 000 ft), respectively.

Also shown in the figure are the calculated curves and lateral "cutoff" distances with current theory being used and a standard atmosphere with no wind being assumed. Again good agreement is noted to exist between theory and experiment.

In addition, both theory and experiment show that the boom overpressures are highest on the ground track and decrease with increasing lateral distance, until the shock waves are refracted at the lateral cutoff (see ref.

34).

Comparisons of the calculated and measured lateral extent of the sonic-boom pat- terns as a function of airplane altitude and Mach number for steady flight in a standard atmosphere are given in figure II-64. The data points represent averages of a number of measurements involving various airplanes.

The widths of the sonic-boom patterns on the ground increase with increasing altitude and Mach number.

For example, at an alti- D(EXPERIMENT M THEORY 1.2 - O 1.5 -- 0 Ap 2.0 . . 0 3.0-- A ~ j " 6.0--- I -d--4 I d M 6.0 100 - --- 3.0

80 . . -- 2.o

1.5 - EXTENT OF ..- AUDIBLE BOOM, d, 40 - 1.2 n. mi.

, I , I I , l 0 20 40 60 80 100 120 x 103 ft I I I I I I I 18 24 30 36 km 0 6 12 ALTITUDE Figure 11-64.- Width of audible sonic boom on the ground.

tude of 6.1 km (20 000 ft) and M = 1.5 the total width of the pattern is 20 n. mi.. At 18.4 km (60 000 ft) and M = 2.0 the pattern width is about 60 n. mi.. However, as is illustrated by the two sketches at the top of the figure, supersonic flights at low altitude result in narrow width patterns having higher overpressures, whereas at higher altitudes the pattern widths are much broader but with lower ground overpressures. Good agree- ment is noted between measured and calculated values. The hypersonic airplanes will operate at altitudes and Mach numbers beyond the current experience. However, there is no reason to believe that theory would not provide reasonable estimates of the pattern width for this flight regime.

Sonic Boom Minimization The discussions thus far have related to the nominal ontrack and lateral sonic-boom pressure distributions, and it has been shown that sonic-boom effects are minimized through increased distance between the airplane and the ground. Minimizing the sonic booms through airplane design modifications have also been investigated (refs. 34 and 35). The basic approaches that have been considered are illustrated in the sketches of figure II-65. Sonic-boom minimization can be achieved through a reduction in the over- pressure or an increase in the signature rise time, each of these parameters being sig- nificant with regard to human and structural response (ref.

36). As illustrated in the REDUCED OVERPRESSURE INCREASED RISE TIME SIZE SHAPE LENGTH AIRSTREAM ALTERATION I I ~I CALCULATED SIGNATURES Figure II-65.- Illustrations of various concepts employed for obtaining sonic-boom minimization.

sketches on the left of figure 11-65, reduced overpressures can be obtained by reducing the size of the airplane (i. e., low airplane weights) or by proper shaping of the airplane geometry to provide a modified signature (a flat top signature). These two approaches have been given much consideration in the past, and reduction in bow-wave overpres- sures of the order of about 30 percent seems obtainable.

More recent minimization techniques involving increasing rise times have been in- vestigated (ref. 37).

If the rise time of the signature could be increased to the point where a sine wave would result instead of an N-wave, the sine-wave pressure signature would not be audible to an outdoor observer although building responses would still re- sult.

In order to obtain finite rise times of the order of 10 to 15 milliseconds (msec), the airplane length would have to be increased by at least a factor of 3 (to about 310 m (1000 ft)) over the greatest length now being considered.

Another means would be to alter the airstream so that the same effects associated with the increased length are ob- tained.

This could be accomplished by the addition of heat or other forms of energy.

Recent studies (refs. 34 and 37) regarding the airstream alteration or "phantom body" concept suggest that large amounts of heat or energy are required (at least the equivalent of the output from four more propulsion engines) to obtain finite rise-time signatures.

The performance penalties associated with the previous approaches and their overall feasibility remain to be evaluated.

Effects of the Atmosphere The effect of the atmosphere on sonic-boom propagation is to cause variations in the measured signatures as illustrated by the data of figure II-66 (see ref. 38). The varia- tions in the bow-wave overpressures as measured along the instrumented linear array of 2438 m (8000 ft) are shown for an airplane in steady level flight. The line represents the nominal overpressure calculated for a standard atmosphere with no wind. Overpressure variation along the ground track is cyclic in nature. The high overpressures are usually N/rn 0 M o- r 4000 60 I lb/ft2 N 0 03 0 0 0 0 020 0 150 '2 0 8000 1800 Figure II-66. - Sonic-boom overpressures as a function of distance along the ground track for a fighter airplane in steady flight at a Mach number of 1.3 and an altitude of 9.3 km 30 500 ft) along with sample signatures.

associated with peaked signatures, whereas the lower pressures are associated with rounded signatures.

These signature distortions are attributed to the lower layers of the atmosphere. The phenomena are thus statistical in nature, and they occur either as a function of time or distance.

A summary of the variations of the ontrack overpressures for steady level flight re- sulting from the atmosphere is given in figure II-67. This statistical analysis comprises most of the planned sonic-boom experiments that have been conducted in this country.

Data are included for a wide range of airplanes, a Mach number range from 1.2 to 3.0, and an altitude range from 3.04 km (10 000 ft) to above 21.4 km (70 000 ft). A total of 12 406 data samples have resulted from 1625 supersonic flights. Although the number of data samples would be considered small from a statistical standpoint, they are large from a logistics standpoint.

Plotted in figure II-67 is a relative cumulative frequency distribution and histogram for ontrack measurements showing the probability of equaling or exceeding the ratio of .9999- 0 o NO. FLIGHTS 625 99 0 DATA SAMPLES 12406 .90 PROBABILITY .10 .50 - .08 RELATIVE .06 .10 - PROBABILITY .04

.02-

.01 0 ! 2 3 4 .001 Pcolc .0001 _meas/ .2 .5 2 APmeas/ Pcalc Figure 11-67.

- Probability of exceeding given values of measured to calculated sonic- boom ground overpressures along the airplane flight track.

the measured overpressure to the calculated or nominal overpressure for steady flight in a standard atmosphere.

For this type of presentation, all the data would fall in a straight line if the logarithm of the data fitted a normal distribution.

Rounded signatures of the wave form shown in the figure are usually associated with overpressure ratios less than 1.

Nominal or N-wave signatures are observed on the average, and peaked signatures of higher overpressures are observed usually at ratios greater than 1. The data of figure II-67 indicate that variation in the sonic-boom signatures as a result of the effects of the atmosphere can be expected during routine operations.

Effects of Airplane Maneuvers Sonic-boom enhancement can result from various airplane maneuvers (ref.

32). In figure II-68 are illustrated three types of maneuvers which could result in pressure buildups at ground level (a longitudinal acceleration, a 900 turn, and a pushover maneu- ver). In each maneuver, pressure buildups occur in the localized regions suggested by the shaded areas shown in the sketches of figure II-68.

It should be pointed out that al- though the airplane and shock waves are moving, the areas on the ground in which pres- sure buildups occur are fixed and do not move with the airplane.

The pressure buildups in these focus areas are a function of the type of maneuver and acceleration involved.

Operationally, pressure buildups will always result for the longitudinal maneuver when ,LONGITUDINAL , /PUSHOVER S ACCELERATION 900 TURN 1 Figure II-68. - Areas on the ground exposed to focused sonic booms resulting from three different airplane maneuvers.

LONGITUDINAL PUSHOVERS, TURNS ACCELERATIONS APfocu 6 s Ap steady 4 .

0 I 0 1 ACCELERATION, g units Figure II-69. - Effect of airplane acceleration on sonic-boom overpressure buildups in maneuvering flight.

the airplane accelerates from subsonic to supersonic speeds.

The effects can be mini- mized by reducing acceleration rates. The pressure buildup areas associated with turns and pushover maneuvers can be minimized or avoided by reducing acceleration or by simply avoiding the maneuver.

The overpressure buildups, or focus factors, in these maneuvers are shown as a function of airplane acceleration in figure II-69. The information contained in this figure is a result of flight experiments conducted in this country (refs.

32 and 39) and in France (refs. 40 to 42).

The data points shown represent the highest levels measured thus far for longitudinal accelerations and turning maneuvers.

The hatched boundary, therefore, would represent the current upper bound.

At the lower values of acceleration, which are usually associated with longitudinal accelerations, the overpressure in the focus is of the order of about 2 to 5 times the nominal overpressure for steady level flight at the same altitude and Mach number.

For a turn maneuver involving a 2 g acceleration, a focus factor of up to 9 has been measured.

Low Supersonic Operations The discussions thus far have been concerned with airplane operations in which the shock waves extend to the ground and are reflected from the ground as illustrated in the upper right sketch of figure II-70. Now consider low supersonic operations where the shock waves extend down toward but do not intersect the ground, as suggested by the sketch in the upper left of the figure.

The range of Mach numbers and altitudes over which cutoff Mach number operations can be performed is shown in figure II-70, with steady level flight in a standard atmosphere with no wind being assumed (see ref.

43).

Flights at Mach numbers to the left of the hatched curve will result in no booms reaching the ground, whereas flights at Mach numbers to the right of the curve will result in booms reaching the ground.

In figure II-70, the highest speed at which the airplane could operate in a standard atmosphere without producing booms at the ground is about M = 1.

15. In the real atmosphere, variations in sound speed gradient do exist because of temperature and winds.

For example, headwinds at altitude and higher temperatures at the surface would increase Mcutoff; conversely, tailwinds at altitude and colder tem- peratures at the surface would decrease Mcutoff.

The practical range of Mcutoff for a fairly wide range of atmosphere (ref. 44) is shown to vary from about 1.0 to about 1.3.

A description of the cutoff phenomenon may be obtained by examining the ray-shock diagram shown in figure II-71.

The top sketch relates to the operating conditions in which the airplane velocity over the ground V is greater than the sound speed at the ground ag. The airplane is shown moving in a direction from right to left and, for sim- plicity, only the bow wave is shown.

The shock wave extends to the ground and is re- flected upward.

As the disturbances that form this shock wave are emitted from the ALTITUDE km ft / ,77, Mcutoff 18 60 000 - BOOMS OBSERVED - 40000 - NO BOOMS , OBSERVED ' 6 - 20000- 0 ' 0- 1.2 1.4 1.0 M Figure 11-70.- Combinations of airplane Mach number and altitude for complete sonic-boom cutoff for steady level flight in a standard atmosphere with no wind.

V >a SHOCK RAYS .

REFLECTED N , -s ', , BOW '. -,. - ,... SHOCK .: INCIDENT "N -. RAYS REFRACTED BOW SHOC REFRACTED RAYS LINE TANGENT Figure 11I-71.- Ray-shock diagram for airplane operating at Mach numbers above and below the sonic-boom cutoff Mach number.

airplane, they travel toward the ground along ray paths indicated by the solid lines.

These rays intersect the ground and are reflected upward, as illustrated by the dashed lines. Any number of such rays can be drawn from the airplane at different times along the flight path. Two consecutive rays are essentially parallel and tend to converge only slightly as they approach the ground.

The bottom sketch of figure II-71 relates to the conditions when the airplane speed over the ground V is less than the sound speed at the ground ag. These conditions would be associated with an airplane flying below the cutoff Mach number. The rays have a greater curvature than those for the higher airplane speed and are totally re- fracted at an altitude above the ground.

The rays, which have been drawn for various positions of the airplane along the flight track, become tangent to each other at the cutoff altitude. The areas between adjacent rays, as indicated by the shaded region, decrease until they theoretically approach zero where the rays become tangent. The pressures are expected to markedly increase in the regions where the area between the rays is de- creasing.

Thus, a caustic, or line focus, is formed where the rays become tangent at the cutoff altitude.

Recent experiments have been made to define the pressure field associated with the shock-wave extremities for flights at low supersonic Mach numbers (ref. 34). The re- sults shown in figure II-72 are indicative of the sonic-boom phenomenon associated with flights at cutoff Mach number.

The experiments used a 457-m-high (1500-ft) tower with microphones placed at 30-m (100-ft) intervals along the tower and also along the ground -

Vg

<

g

_ BOOM REGION - --- - - 457-m (1500 ft) - _ _ i_- TOWER - CAUSTIC (LINE FOCUS) ACOUSTIC REGION Figure 11-72.- Measured sonic-boom signatures at various heights above the ground for an F-104 in steady level flight at cutoff Mach number. Altitude, 10.26 km (33 700 ft); = M 1.095.

so that the shock patterns terminated within the (see ref. 45). The airplane was flown shock waves, and the dashed lines represent tower heights. The solid lines represent the refracted waves. Also shown in the figure are the measured sonic-boom signatures observed as the shock-wave system moved across the microphones on the that were tower. In this particular experiment, the shock extremity was positioned so that pres- disturbances in the supersonic, sonic, and acoustic regions were measured. At the sure top of the tower, a normal N-wave sonic-boom signature was measured that included the incident bow and tail shocks and their refractions. At midtower, which was the approxi- a U-shape (caustic) signature was measured for mate location of the shock extremity, ob- which the amplitudes were larger than those associated with the N-wave signatures the base of the tower, below the shock ex- served in the supersonic flow region. Near disturb- the signatures are rounded or sinusoidal in shape and suggest acoustic tremity, ances.

signatures such as those shown in figure II-72 for the The bow overpressure from presented in figure II-73. The measure- flight operations near cutoff Mach number are which represent the average values obtained from the ground and tower micro- ments, level flights of the F-104 airplane at 10.26 km (33 700 ft). The phones, are from steady solid curve represents the overpressure variations suggested by theory, and the data somewhat (from about points represent the measurements. Because Mcutoff varied M = 1.09 to 1.12) from flight to flight and over the 3-day test period because of varia- the measured data have been normalized so that the tions in atmospheric conditions, located at a cutoff overpressures associated with the caustic or U-shape signatures are Ap _ Ap

i4,~

N/rn ' 1P LIL

Ib/ft 200 - 150- 3 - o MEASURED -THEORY 100- 2 - 50- 1 eba 50 - 1 - 0- 0 I 0 McutoffRANGE I 1.0 1.1 1.2 1.3 1.4 1.5 M Figure II-73.- Variation of sonic-boom overpresure with airplane Mach number for an F-104 in steady level flight at 10.26 km (33 700 ft).

Mach number of M = 1.1.

At Mach numbers below M = 1.1 very low overpressure levels were measured, and these were associated with the acoustic type of signature shown at the top left of figure II-73. At Mach numbers greater than M = 1. 1, normal N-wave types of signatures were observed from which the pressure increased gradually with increasing Mach num- ber as predicted by theory.

Near M = 1.1, U-shape waveforms were generally ob- served.

The predicted overpressure values would approach infinity because of the pres- ence of the caustic line.

The measured caustic signatures generally indicate overpres- sure enhancement compared with those associated with the higher Mach numbers.

The highest enhancement factor suggested by the data of figure II-73 is 3.

The experiments also suggested that very stable conditions of the atmosphere, especially the lower layers, are required to produce the U-shaped signatures of high overpressures.

Summary of Results The information derived from a number of recent flight and wind tunnel studies is summarized in figure II-74, in which are shown the variations with Mach number of the nominal overpressures for cruising flight in the low supersonic, supersonic, and hyper- sonic operating regimes.

The hatched area represents the amount of boom minimization that may be attainable through airplane design.

Above a Mach number of about 2, increasing Mach number results in decreasing LOW SUPERSONIC A p _ 2 2 N/m Ib/ft SUPERSONIC 150 - HYPERSONIC - - I I I I 3 5 7 9 M Figure II-74.- Variation of sonic-boom overpressure along ground track with airplane cruise flight speed.

overpressure.

The decrease in overpressure results primarily from the increased alti- tudes required for flight efficiency and range as the speed increases. In attaining cruis- ing Mach numbers, however, both the supersonic and hypersonic vehicles would pass through the low altitude Mach number region, where higher overpressures would be ex- perienced. As the cruise Mach number is increased, sonic-boom minimization through airplane design is shown to be less effective.

Perturbations on the nominal values can be expected as a result of routine opera- tions in the atmosphere. In addition, pressure enhancement resulting from maneuvering flight would also increase the nominal overpressure values. Certain of the overpressure buildups resulting from maneuvering flight are unavoidable; however, others are avoid- able or can be minimal. Studies, thus far, suggest that the atmospheric and maneuver pressure enhancements are not additive. In conducting low supersonic operations in- tended to prevent shock waves from extending to the ground, great care must be exer- cised to match airplane operations to local weather conditions to avoid the sudden onset of large amplitude transonic booms.

Planned Research Planned research relates directly to developing improved understanding of the mini- mization of nominal boom intensity and the prediction of occasional pressure enhance- ments or focusing due to accelerated flight and to atmospheric inhomogeneities.

The reduction of the nominal sonic boom intensity by configuration shaping is under continu- ing study and involves the use of analytical computer studies and complementary wind tunnel tests. The development of improved prediction methods for focused booms in- volves analytical studies and related precision measurements in focus regions by means of controlled acceleration rocket sled tests.

POWERED LIFT AIRCRAFT Commercial powered lift aircraft will be needed to reduce congestion around our major cities. They will probably be operated from existing and/or new airports, with short runways, in or very close to cities. The cost and complexities of the lift and pro- pulsion systems will probably limit their use to short-haul applications.

The powered lift effort presently is being concentrated on three principal aircraft types: (1) The augmentor wing (AW) (2) The externally blown flap (EBF) with engines located under the wing (UTW) (3) The EBF with engines located over the wing (OTW) CTOL AUGMENTOR WING UNDER-THE-WING EBF OVER-THE-WING EBF cs-64891 Figure II-75. - Short-haul and CTOL aircraft These aircraft concepts are shown in figure 11-75 compared to a conventional CTOL air- craft.

The EBF powered lift and propulsion systems are shown in figure II-76 and the AW system in figure II-77, together with the significant noise sources.

In all three concepts the exhaust from the propulsion system is directed onto or through the wing airfoil. Increased lift is thus obtained by the downward turning of the exhaust jet and by the increased flow induced around the wing by the jet.

Because powered lift aircraft will provide service out of heavily congested areas, the environmental specifications that will be imposed on them will be very severe. Ex- tremely quiet and pollution-free operation will be demanded. Noise specifications for these systems have not been established; however, NASA has selected 95 EPNdB on a 500-foot sideline as a preliminary goal toward which to direct their research efforts (refs. 46 and 47).

Figure II-78 illustrates a 100 PNdB noise footprint for the DC-10-10 wide body jet compared with the footprint for a 95 EPNdB powered lift aircraft. Dramatic reduc- tions in area are achieved by the 95 EPNdB aircraft.

The severity of the noise goal can readily be appreciated by referring to figure 11-79. This figure compares the present FAA FAR-36 noise requirement for conventional takeoff and landing (CTOL) aircraft, extrapolated to a 500-ft sideline measuring point, with the powered lift goal and the NASA quiet engine A. Note that the goal is approximately 30 PNdB lower than existing FAN FAN& & L FLAP IMPINGEMENT NOISE '-CORE & NOISE JET NOISE JET NOISE "UNDER-THE-WING" CONCEPT r JET & MACHINERY ~/ NOISE SrWING SCRUBBING FAN NOISE" NOISE "OVER-THE WING" CONCEPT CS-63126 Figure II-76. - Blown flap powered STOL system.

FAN -NOISE-, AUGMENTOR MET & CORE NOISE Figure II-77. - Augmentor wing propulsion system.

CS-63128 4- 2 - DC-IO-IO 2 - " ,SHORT-HAUL D-01 SIDELINE DISTANCE, -2 1000 FT -4 - -6- -8 I I 1 I I 1 1 11I1 1I I -14 -12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 DISTANCE FROM START OF RUNWAY, 1000 FT Figure II-78. - Estimated 100 PNdB noise footprints for a DC-10-10 and for short-haul airplane capable of 95 PNdB on 500 ft sideline. Mission 500 nautical miles.

SIDELINE TO 500-FT 130 *EXTRAPOLATED FAR-36* 120- NASA EPNdB QUIET AT 500 FT 110 - ENGINE A* SIDELINE 100 - I STOL GOAL 90- II-79. - Sideline noise Figure comparisons. Thrust, 90 000 lb.

regulations and 13 PNdB less than was achieved with quiet engine A. The effort re- quired to achieve this goal is made more difficult because powered lift systems generate an additional noise source as compared to CTOL systems. The externally blown flap systems have flap impingement and scrubbing noise which results from the flow of engine airflow over and against wing and flap surfaces.

The augmentor wing lift system has augmentor noise from the discharge of a high pressure jet into the flap system.

Addi- tionally, when comparing powered lift aircraft noise to CTOL noise, one finds that there is an approximate 3-dB penalty associated with powered lift systems due to the fact that their thrust-weight ratios are roughly twice those of CTOL systems.

This difficult noise goal now dominates the design of AW and EBF aircraft as is shown in the following sections.

AUGMENTOR WING NOISE The AW and its propulsion system together with the major noise sources were shown in figure II-77. The fan exhaust of the turbofan engine is ducted upwards through the pylon to the rear of the wing. At this point, the exhaust air is distributed span-wise and then discharged aft between two parallel flap segments.

The flaps are set at different angles for takeoff and landing. In the cruise mode, the flaps retract and the fan air is ejected through a separate nozzle.

Ejecting the wing air between the two flaps creates an ejector action which augments the lift but, more importantly, augments the forward thrust by an amount that more than makes up for any duct pressure losses. Current AW designs require fan pressure ratios of about three to meet wing volume requirements and noise restrictions (refs. 48 and 49).

Reduction of fan machinery, core, and core jet noise represent problems similar to those described under the SUBSONIC TRANSPORT AIRCRAFT section except that the more stringent powered lift aircraft noise goal further compounds the problem.

The AW fan is a multistage high pressure ratio fan and, as was shown earlier, is very noisy. The fan noise which radiates forward must be reduced about 30 dB to meet the goal. The sonic inlet previously described has the potential of providing this sup- pression. The fan noise radiated rearward is contained within the takeoff and distributor ducts to the wing and can be suppressed with acoustic wall treatment. The internal core noise originating from the turbine, combustor, etc., may be reduced using techniques described earlier except the lower levels desired makes the problem even more diffi- cult. Core jet noise may be reduced to the required level by extracting more energy from the core flow to reduce its velocity to about 800 ft/sec. The resulting engine cycle will have a fan pressure ratio of about 3 and a bypass ratio about 2. 5.

Without suppression and using a single rectangular slot as a wing exhaust, the high pressure fan exhaust flow discharges from the wing nozzle at 1200 to 1300 ft/sec and produces noise levels in excess of 115 PNdB at a 500-ft sideline if incorporated on a 100-passenger aircraft (refs. 50 to 53).

Since noise levels are unacceptable, NASA conducted a 2-year program with The Boeing Company to improve the augmentor wing performance to reduce and suppress the noise. The program included static thrust and noise tests of nozzles and nozzle-flap combinations, wind tunnel tests, and design integration studies.

Instead of the single rectangular slot nozzle employed in the early programs, Boeing studied the use of a wing nozzle design that had many small lobe nozzles.

The use of multiple nozzles increased the ejector action of the flaps and hence the thrust augmentation of the system. More importantly, the several small, multiple nozzles raised the frequency of the jet noise which allow the practical use of acoustic treatment on the interior flap surfaces. (Acoustic treatment thickness is related to the noise wavelength.) The improved ejector action entrained more ambient air which mixed with and reduced the jet velocity exiting from the flap system to below 500 ft/sec.

Thus, the wing nozzle-jet noise is effectively contained and suppressed within the flap system. A sketch of a full augmentor installation is shown in figure II-80 including details of the ducting and the multiple lobe nozzles (ref. 52).

The noise reductions obtained by the various steps taken in this development are shown in figure II-81 in terms of the 500-ft sideline noise level as a function of the nozzle pressure ratio. The term "soft flap" refers to the use of acoustic treatment.

The noise level of the final optimized system is seen to be below the 95 PNdB goal for all pressure ratios of interest.

Finally, the unsuppressed and suppressed noise levels of all the component noise sources are shown in figure II-82 for a typical short-haul transport.

All sources have been suppressed sufficiently to attain the 95 PNdB side- line goal for the total aircraft (ref. 53).

REAR SPAR-I SECONDARY N CS-64892 SECTION A-A Figure II-80. - Augmentor duct and nozzle details.

120- SLOT 10SLOT IN HARD FLAP 500- FT 70-LOBE HARD FLAP SIDELINE 172-LOBE SOFT FLAP PNLB 100 D PNdB OPTIMIZED SYSTEM 90- 801 I 1 I I I 2.0 2.5 3.0 NOZZLE PRESSURE RATIO Figure II-81.

- Augmentor jet-flap noise reduction.

Four 20 000-lb SLS thrust engines; takeoff, 100 KTS.

- UNSUPPRESSED 120 - --- SUPPRESSED 110 - 500 FT MAX SIDELINE NOISE. 100 - PNdB GOAL 8D FAN CORE JET WING INLET cs-s4774 Figure II-82.

- Augmentor wing component noise sources.

4 Engines, 80 knots; 60 000-lb thrust; 3.0 fan pressure ratio.

(approximately 1/6).

Boeing contract was at small scale All the testing under the the procurement for on the AW systems include plans for future research Current spanwise section of a full scale, partial Lewis Research Center acoustic testing at the nozzle, flap, test configuration will have the latest and most advanced of a wing. This underway with re- planning is currently designs. Preliminary and acoustic treatment tunnel. This system system in a large wind to testing a semispan augmentor spect ducting and valving.

use real engines and have realistic would to aircraft can be designed indicated that an AW short-haul Research to date has noise goals.

meet the stringent BLOWN FLAP NOISE EXTERNALLY propulsion sys- is achieved by locating the In the EBF systems, lift augmentation system (or are directed over) the wing-flap the exhaust gasses impinge on tems so that about by the reaction to the Lift augmentation comes and then deflected downward.

pressure distribution on the also by inducing a favorable downward jet deflection and wing-flap system.

blowing systems consists of the by both the UTW and OTW The noise generated noise, noise of engine itself, fan machinery, core and jet noise, plus wing scrubbing an fig. II-76). The OTW installation has and trailing edge noise (see and flap leading that it attaches to the is used to position the jet so additional noise source if a deflector in that provides a major acoustic benefit However, the OTW installation wing surface.

in flight.

a substantial amount of the noise the wing-flap system shields by the jet-wing interactions dominates the design Reduction of the noise generated be shown.

and engine acoustic design as will of the EBF aircraft of jet noise section showed that, if a noise level The data in the subsonic aircraft than 850 ft/sec is required for both the PNdB is to be achieved, a jet velocity less been found that still flow. For the EBF powered lift systems, it has fan flow and core are required to reduce the powered lift source further reductions of exhaust velocities to the desired level. Studies have noise (e.g., flap impingement and wing scrubbing) is with high the most economical means of achieving these low velocities indicated that and low fan pressure ratios (FPR) to bypass ratio engines to achieve low core velocities bypass ratios velocities. Engines with FPR of 1. 25 to 1. 30 and achieve low fan exhaust for EBF UTW. Because of the wing shielding effects of 12 to 20 appear optimum of 10 to 16 appear optimum for slightly higher FPR's of 1.30 to 1. 35 and bypass ratios EBF OTW installations.

still require signi- low pressure ratio fans are quieter than CTOL fans but These the*noise goal. The use of conventional acoustic sup- ficant noise suppression to meet earlier, should provide pression in conjunction with the high velocity inlets, described this suppression.

Core noise suppression presents a similar, equally difficult, problem to that en- countered with the AW.

The major new acoustic problem, jet-flap noise, is now discussed in detail.

Research and development activity in the UTW (refs. 55 to 57) and OTW area has consisted largely of small-scale cold gas tests of various exhaust nozzle and wing sec- tions plus two full-scale engine test programs. The large-scale tests were conducted with a CF700 engine and F-111 wing and flap and with the highly suppressed TF-34 en- gine, discussed earlier, with a triple slotted EBF wing section.

Plotted in figure II-83 are the results from large-scale cold gas tests which com- pare the noise patterns for both the upper surface and lower surface blowing powered lift systems (refs.

58 to 61). The noise directionality and the shielding of the OTW in- stallation are evident. Results of all testing to date show that flap noise can be con- trolled mainly by reducing the impingement velocity of the fan and core air stream. Re- sults from cold gas model tests, shown in figure II-84, indicate that only relatively low exhaust velocities will provide a noise level acceptable for STOL systems (refs. 62 and 63). These low velocities are not typical of today's fan engines and can best be achieved with fan pressure ratios of about 1. 25.

Lower flap impingement velocities can also be obtained with special exhaust nozzles (mixer decayer) which break up the fan stream into smaller jets, thus producing more rapid mixing and large velocity decay over a shorter distance (refs. 64 and 65).

The TF-34 engine and acoustic nacelle were evaluated with various nozzle com- binations and decayers (fig. II-85) with a wing section for both the UTW and OTW con- figurations. Results of this program are summarized in figures II-86 and II-87. Sche- *-- LOWER SURFACE BLOWING SUPPER SURFACE BLOWING PERCEIVED 110 - LEVEL, NOISE NOISE LEVEL, SUPPER SURFACE BLOWING PNdB 90 . LOWER SURFACE BLOWING II o cs-64708 0500 600 700 800 900 1000 CS-64707 NOZZLE EXHAUST VELOCITY, FTISEC Figure II-83. - Comparison of EBF perceived noise level Figure II-84 - Effect of nozzle exhaust velocity on EBF patterns at 500 ft. Nozzle diameter, 13 in. ; wing chord perceived noise level at 500 ft Nozzle diameter, 13 in.

length, 7 ft; flap position, 300-60o; exhaust velocity, wing chord length. 7 ft; flap position, 300 - 600.

ftisec.

MIXER-DECAYER CO-PLANAR DECAYER MIXER-CONICAL CO-PLANAR DECAYER WITH 3-FOOT SHROUD CONICAL-CONICAL CO-PLANAR DECAYER WITH 4-112-FOOT SHROUD CONICAL-DECAYER - BASELINE MIXED FLOW CONFIGURATIONS SEPARATE FLOW CONFIGURATIONS Figure II-85.

- TF-34 exhaust nozzle configurations.

SUPPRESSED ENGINE WITH WING AND FLAPS IN TAKEOFF POSITION * MIXER-DECAYER o SEPARATE FLOW CO-ANNULAR 0 COPLANAR DECAYER * COPLANAR DECAYER WITH SHROUD o MIXER-CONICAL OVER-THE-WING 110 - o SEPARATE FLOW-COANNULAR-NO WING 100- MAXIMUM PNL AT 500-FT FLY-OVER, PNdB (SINGLE 90 ENGINE, NO FORWARD VELOCITY) 70 I I I I I I 3000 4000 5000 6000 7000 8000 9000 CORRECTED THRUST, LB Figure II-86. - TF-34 noise comparison for various exhaust nozzle configurations.

matics of the nozzle configuration tested are shown in figure II-85. For the takeoff flap configuration, it is seen (fig. H-86) that the flyover suppressed engine noise, with no wing, increased approximately 12 PNdB when the fan and core flow from a separate flow coannular nozzle impinged on the flaps. A 12-lobe internal and external mixer-decayer was installed to reduce the impingement velocity. It was found, however, that although the velocity decay achieved was more than required, 8 PNdB of noise were generated by the mixer-decayer (fig. II-87). Tests of the mixer-decayer with the wing and flap sys- tem produced a noise level increase of 13 PNdB over that from the suppressed engine without wing.

A redesign of the decayer proved effective in reducing mixing noise.

To reduce the mixing noise still further, an acoustic shroud was placed over the decayer.

With this final configuration, it was found possible to limit the UTW noise to an increase of 5 PNdB over that of the suppressed engine without a wing.

All the previous results are shown plotted (fig. II-86) for a flyover noise condition. Measurements were also made to establish the sideline noise with the various test configurations. These data (fig. II-87) resulted in an average sideline noise reduction of 5 PNdB from that of the flyover noise.

This reduction was due to directionality differences in the noise patterns from the flyover to the sideline measuring points.

The sideline noise for the best UTW exhaust configuration was equal to that of the suppressed engine without a wing.

One test was conducted to establish a data point for the noise of an OTW lift system using a TF-34 engine.

The quietest suppressed engine exhaust configuration was used, TAKEOFF THRUST 500 FT SIDELINE 6235 SLS (WITHOUT EJECTOR) 6670 SLS (WITH EJECTOR) 0 - UNSUPPRESSED S-3A 1 - SUPPRESSED - SEPARATE FLOW 2 - SUPPRESSED - MIXERIDECAYER 3 - SUPPRESSED - CO-PLANARIDECAYER 4 - SUPPRESSED - CO-PLANARIDECAYER AND EJECTOR 5 - SUPPRESSED - MIXERICONICAL - OTW ENGINE ALONE ENGINE AND WING 10- 100- MAXIMUM PNLI 95 PNdB L 95 3 WITH JET DEFLECTOR PNdBPLATE 90 89 89 89" WITHOUT JET DE- FLECTOR 0 1 2 3 4 5 1 2 3 4 5 Figure II-87. - TF-34 noise summary.

into a conical fan exhaust duct. A deflec- which was an internal core mixer discharging tor plate was added to the exhaust exit to attach the flow to the wing. A penalty of 7 PNdB resulted from the use of the deflector plate (fig. 11-87).

Figure II-86 shows that the OTW flyover noise is quieter than that obtained with the best UTW exhaust system. The sideline noise was found to be only 1 PNdB higher than that of the suppressed engine by itself.

QCSEE PROGRAM The Quiet, Clean Short-Haul Experimental Engine (QCSEE) Program is a major element in NASA's quiet powered-lift propulsive technology program. It is being under- taken to establish the technology base for very quiet propulsion systems, designed for installation in powered-lift aircraft.

The major objective of the QCSEE Program is to develop and demonstrate the technology required for propulsion systems for quiet, clean, and economically viable commercial EBF powered lift short-haul aircraft. As was shown, quieting the aircraft is the most challenging problem.

The QCSEE Program consists of a definition phase already completed which has defined the engine for the short-haul powered-lift aircraft, and a hardware fabrication and testing phase to provide a demonstration of a quiet, economically feasible experi- mental EBF propulsion system.

The definition phase was initiated with in-house studies in May 1971. Two QCSEE Electric were completed in propulsion system study contracts with Allison and General January 1973 (refs. 66 to 69). In order to strengthen the engine definition phase, the two section were also con- parallel aircraft system study contracts described in the next ducted. The engine contractors, aircraft contractors, and NASA worked closely together to define optimized economically viable quiet propulsion systems for the UTW, OTW, and the augmentor wing powered-lift concepts. Results from these major study efforts and cQmponent technology programs were used to define the optimum propulsion systems for short-haul aircraft.

The component program was conducted in-house and under contract. It is providing a sound technical background for the engine studies and will be used in the future QCSEE design efforts. The programs include (1) EBF Fan Aero Development (2) EBF Fan Acoustics (3) High Mach Number Inlet Development (4) Acoustic Suppression Development (5) EBF Flap Noise Reduction These programs were summarized earlier.

Details are, or will be, available in forthcoming NASA reports.

The completed program effort has indicated that the noise goal can be met by unique design approaches, but more extensive testing of large-scale propulsion and powered lift systems are essential before the full extent of the cost is determined.

The effort has shown the following: (1) A variable pitch fan (1. 25 to 1. 30 pressure ratio) propulsion system on an EBF- UTW aircraft might meet the noise goal without excessive cost penalties. Figure II-88 is a sketch of this engine.

(2) The AW and EBF-OTW systems also have a good potential for meeting the noise goal with more conventional propulsion systems.

The installation of these propulsion systems in the aircraft, however, presents many unique problems.

,.-FAN DUCT ACOUSTIC RINGS CRUISE NOZZLE POSITION .. W \''-TAKEOFF NOZZLE POSITION INLET WALL I ACOUSTIC ACCESSORY ' -REVERSE PITCH POSITION TREATMENT-' GEARBOX -FAN DUCT WALL 'INLET DUCT ACOUSTIC TREATMENT ACOUSTIC RINGS cs-64769 Figure 11-88.

- Variable pitch fan engine.

EBF aircraft (UTW).

CALENDAR YEAR 1972 1973 1974 1975 1976 1977 STOL QCSEE STUDIES 6 6 STOL AIRCRAFT SYSTEM STUDIES -- SUPPORTING COMPONENT DEVELOPMENT ISSUE RFP CONTRACT AWARD DESIGN - COMPONENT AND PROPULSION SYSTEM COMPONENT DEVELOPMENT ENGINE AND PROPULSION SYSTEM DEVELOPMENT: FAB DEVELOPMENT - ENGINES ENGINE AND PROPULSION SYSTEM TESTS PROPULSION SYSTEM DELIVERY TO LeRC PROPULSION SYSTEM TESTING Figure II-89. - QCSEE program plan.

• 97 on the development of the the QCSEE Program will concentrate The second phase of installations.

for EBF OTW and UTW fan engine and the "quiet" nacelles variable pitch using current fiscal guideline for the QCSEE Program Figure II-89 depicts the plan restraints.

SYSTEM STUDIES NASA SHORT-HAUL AIRCRAFT let to the Douglas Aircraft QCSEE studies, contracts were In parallel with the of Quiet Turbofan Company to perform a "Study and the Lockheed Aircraft Company " These contracts were managed by the STOL Aircraft for Short-Haul Transportation.

Both contractors use and Mission Analysis Division at Ames.

Advanced Concepts conducted as subcontractors. The contractors inter- and intra-state airline companies analysis of a large number aircraft (refs. 48, 49, and 70) design an in-depth parametric and levels of performance; screened the large of candidate aircraft concepts, sizes, representative of the against a parametric transportation system matrix of designs six STOL aircraft. The aircraft companies national short-haul market; and designed powered-lift aircraft. Both the the optimum propulsion systems for these also defined during aircraft manufacturers maintained close liaison engine manufacturers and the a sound basis for further R&D the studies. The studies served to guide and to provide on viable quiet short-haul air transport systems.

These included augmentor Several powered-lift aircraft systems were analyzed.

flap with engines mounted both OTW and wing, externally (EBF) and internally blown to powered- mechanical flap aircraft were analyzed for comparison UTW. In addition, A substantial part of the early study effort lift systems for convention field lengths.

candidates submitted by the QCSEE engine was devoted to evaluation of various engine as part of the related contract.

study contractors it was concluded that: From the studies, pitch fan was optimum for meeting the (1) A low fan pressure ratio (1. 25), variable takeoff field length.

95 EPNdB noise goal for EBF UTW aircraft with a 2000-ft the noise fan pressure ratio propulsion system could reduce cost if (2) A higher 5 dB. The effect of this relaxation on the footprint area is shown goal is relaxed 3 to in figure II-90.

The aircraft economic systems analysis showed that for field lengths below 3000 ft, competitive costwise although there are differences in all powered lift concepts are noise footprints as shown in figure 11-91. For field lengths greater than 3000 ft, the conventional mechanical flap (MF) systems start showing economic benefits. However, noise goal is to be met.

fan pressure ratios are limited to less than 1. 35 if a 95 EPNdB FOOTPRINT AREA, ACRES 95 EPNdB 85 EPNdB 98 EPNdB AIRCRAFT 340 2550 95 EPNdB AIRCRAFT 203 1520 5 - EPNdB 0 - 98 EPNdB AIRCRAFT -5i I 95 EPNdB AIRCRAFT -5 I III -10 -5 0 5 10 15 DISTANCE, 1000 FT Figure I-90. - Footprint comparison of 98 and 95 EPNdB aircraft EBF-UTW 2000 ft field length.

125 Passenger aircraft FOOTPRINT AREA, ACRES 95 EPNdB 90EPNdB AW 213 OTW 205 431 UTW 203 531 EPNdB AW ENGINE REVERSER 2 AW 90 117 EPNdB AT 500 FT 0-- -2 -.

OTW UTW 4 0 4 8 12 16 DISTANCE, 1000 FT Figure II-91. - Footprint comparison of AW, 01W, and UTW aircraft.

95 EPNdB aircraft, 2000 ft field length, and 125 passenger aircraft ROTORCRAFT The two main sources of noise from the rotorcraft are the engines and rotors. The engines may be of the reciprocating, rotary, or turbine types, and their noise charac- teristics can thus vary widely. The noise source characteristics and the noise control technology approaches for gas turbine engines for rotorcraft are similar to those pre- viously discussed for other types of flight vehicles and are thus not included herein.

This section includes discussions of reciprocating and rotary engine noise sources and rotor noise sources.

ENGINE NOISE The dominant noise source for reciprocating and rotary engine powered helicopters is usually the engine exhaust if no muffler system is installed.

Muffler design methods that rely greatly on empiricism are available for providing significant exhaust noise reduction. Types of reactive devices that will accomplish this task are shown schematically in figure II-92 together with their transmission loss char- acteristics. Practical muffler systems consist of combinations of such devices. Al- though simple theory provides a guide for experimentally combining these devices into a workable muffler system, it has not provided realistic quantitative predictions of com- HELMOLTZ QUARTER WAVE EXPANSION CHAMBER RESONATOR RESONATOR WITH BRANCHES co L

L

___

, s 2

siP S , p SrPt P --- -- LP S1 2 P t T.L., dB f c c o 4L 2n IV - FREQUENCY FREQUENCY FREQUENCY Figure II-92.

- Schematic diagrams of the various types of reactive mufflers and their transmission loss characteristics. (From ref. 71.)

plex muffler system performance. At Langley, considerable effort has been devoted to developing a computer program that makes use of an improved theory for predicting the performance of combinations of expansion chamber type mufflers as shown in the sketch of the figure. In addition this computer program incorporates an optimization procedure that allows for some variation of muffler configuration to meet a specified minimum transmission loss characteristic.

The performance of a three-stage expansion chamber muffler for a helicopter engine is shown in figure II-93.

In this figure the unmuffled helicopter noise spectrum is com- pared with the spectrum obtained with the muffler system installed.

Clearly the spec- trum for the modified aircraft is now dominated by rotor noise.

It is anticipated that the incorporation of the capability of handling other reactive devices as well as absorptive elements into a computer program will significantly en- hance the practical usefulness of this analytical approach to muffler design.

This pos- sibility would require further basic research effort to determine the effect of operating parameters such as flow and temperature gradients on resonator performance.

Also, further data would be needed on source impedance effects to predict total exhaust noise reduction.

To date, only a limited amount of experience is available regarding applications of exhaust mufflers to helicopters.

It can be anticipated, however, that the normal payload will be reduced somewhat because of the additional weight of muffler, manifolding, etc., and there may be some relatively small power loss due to increased backpressure on the engine.

STANDARD MODIFIED 3f I9

SOUND 80 I

. h ,1 5fzl, , JA V , PRESSURE FREQUENCY, Hz J I I I J1 LEVELtion and with modification to include exhaust muffler.

From ref. 72.

0 80 160 240 320 400 480 0 80 160 560 640 240 320 400 480 560 640 FREQUENCY, Hz dB 6f Figure II-93. -Comparison of noise radiation from helicopter in standard configura- Figure II-93.

- Comparison of noise radiation from helicopter in standard configura- tion and with modification to include exhaust muffler.

(From ref. 7Z.)

10o1 ROTOR NOISE engines with exhaust In the cases where either turbine engines or reciprocating mufflers are used, the rotor noise may be the dominant noise component for helicopters.

The major problems associated with helicopter blade noise reduction are (1) validation and expansion of existing prediction theories, (2) more exact understanding of the gen- eration mechanisms of each noise category, and (3) reduction of the acoustic effect of the airflow mechanisms with as little performance penalty as possible.

Figure II-94 presents estimated relative perceived noise levels associated with helicopter rotor operation at various tip speeds and for a given thrust, along with indi- cations of the significant sources of noise in the various operating regimes. With re- gard to steady loads, theories are available and have been validated for predicting the lower order harmonics. These predictions are appropriate for long-range detection and fuselage and wing structural response problems. They give results of acceptable engineering accuracy. With regard to the unsteady (periodic and nonperiodic) aerody- namic load sources, however, the available prediction theories have essentially not been validated. Figure II-95 illustrates the importance of the unsteady periodic aerody- namic loads in predicting periodic rotor noise.

High priority research items are the validation of the theoretical concept of fluc- tuating loads and the definition of operating conditions over which it is useful, and the identification of practical means of minimizing rotor noise for particular applications.

High-frequency blade surface pressure data are required before any significant progress can be made toward understanding aerodynamically generated sound. The occurrence of stall, the degree of compressibility, and the severity of blade-wake in- l/// MEASURED DATA --- FLUCTUATING LOADS THEORY NOISE -i"- STEADY LOADS THEORY 10 PNdB 80 - O LEVEL Hz f.STEADY LOADS RELATIVE r 0 PNL m 60 - S ..... j.P"PERIODIC LOADS OVERALL ........... .NOISE NONPERIODIC LOADS LEVEL, dB I I I I 100 150 200 250 300 ROTOR TIP SPEED, MISEC I I I I I I 400 600 800 1000 0 1 2 3 4 ROTOR TIP SPEED, FTISEC HARMONIC NUMBER Figure II-95. - Comparison of measured Figure II-94. - Schematic diagram relating perceived noise levels of unducted rotors as a function of and calculated helicopter rotor ro- tip speed for a constant rotor thrust. (See ref. 73.) tational noise for several harmonic numbers. (See ref. 74.)

teractions must be known in connection with three-dimensional transonic and supersonic dynamic airfoil characteristics. The azimuth position at which a rotor blade generates maximum sound must be specified accurately, a requirement which imposes much com- plexity both in analytical treatments and experimental flight-test data and analyses.

Therefore, controlled whirl tower and quiet wind-tunnel testing of rotors with advanced blade aerodynamic and acoustic instrumentation holds the greatest promise of isolating and treating individual aerodynamic sound sources.

Figure II-96 indicates some of the methods being used to measure the aerodynamic inputs for the noise prediction theories. Emphasis in these programs is on the identifi- cation of trends in dynamic loading which can be related to standard lift and drag param- eters.

ROTOR PRESSURE GAGES ,, RECEIVER-

) ) ) -MICROPHONE

'-TELEMETER COMPRESSOR AIRFOIL , HOT WIRES , R'LpRESSURE GAGES - PRESSURE GAGES Figure II-96.

- Three example measurement situations for obtaining fluctuating blade loads information.

A recent collaborative effort by the DOD and NASA to use existing technology to reduce helicopter external noise resulted in the Quiet Helicopter Program. Three dif- ferent types of helicopters were modified in this program and were evaluated from an acoustic and performance standpoint.

Modifications to the helicopters included the ad- dition of blades to both the main and tail rotors, reduction of main and tail rotor rota- tional speeds, different rotor blade tip shapes, engine inlet and exhaust noise suppres- sion, and miscellaneous noise control work directed at the power train.

Differing amounts of noise reduction were obtained with the various helicopters. Figure II-97 shows the helicopter which demonstrated the largest overall ontrack noise reductions (approx. 14 dB) of the three tested (see ref.

75).

In certain helicopter flight conditions, the predominate noise source is rotor blade Figure II-97. - The Hughes OH-6A helicopter modifications to reduce external noise.

NORMAL TIP VORTEX CORE WITH TIP MODIFICATION AIR SOURCE WITH AIR MASS INJECTION Figure II-98. - Schematic illustrations of two methods of blade tip vortex dissipation for noise reduction purposes.

"bang." This impulsive noise is caused by localized impulsive aerodynamic loading which occurs periodically.

The low-speed periodic loading occurs when a blade passes close to a wake vortex filament and experiences a change in angle of attack.

Several methods are proposed for reducing impulsive noise due to blade/vortex interactions; these are illustrated in figure II-98.

The first method involves the passive diffusion of the tip vortex by an ogee tip, the second method is an active system which affects the vortex strength by air mass injection.

PLANNED RESEARCH Programs planned for future helicopter rotor research are (1) measurements of fluctuating pressure patterns on a static airfoil at various angles of attack for correla- tion with radiated noise, (2) studies of the effects of tip mass injection to alleviate noise associated with the tip vortex structure, and (3) studies of the effects on radiated noise of nonuniform inflow to the rotor disk. In all rotor noise tests the acoustic results will be correlated with performance data. The trend toward lower tip speed rotors for noise control will probably result in relatively heavier blades and hubs.

GENERAL AVIATION AIRCRAFT General aviation aircraft constitute the fastest growing segment of the aircraft in- dustry. They are operated in both urban and rural areas and usually at relatively low altitudes.

The main noise sources are the engines and the propellers. Internal com- bustion engines provide the power for the large majority of general aviation aircraft and, because of economic considerations, this situation will probably continue for a long time.

Turboshaft engines are preferable from a noise reduction standpoint but are markedly more costly.

The procedures for internal combustion and gas turbine engine noise con- trol have already been discussed in the previous sections and are thus not included herein. This section includes discussions of propeller noise and the general technology associated with quiet general aviation vehicle design.

QUIET VEHICLES The development and demonstration of relatively quiet single engine propeller driven aircraft has extended over approximately a 25-year period and several specific flight test projects have evolved (see refs.

76 to 80). The resulting quieted aircraft have involved the use of engine exhaust mufflers, modified propellers and gear boxes for reduced noise, and in some cases provisions for intake and accessory noise control.

These flight vehicle projects have demonstrated the basic technology required for de- velopment of quiet general aviation vehicles.

QUIET PROPELLERS Quiet propellers generally involve lower tip speed operation and multiblade designs.

Theories are available for predicting the noise both near the propeller as it may affect the fuselage and its occupants and in the radiation field.

Prediction of the noise due to steady loads on the blades has been well established and confirmed by experiments (see refs. 81 and 82). Methods of predicting noise due to the unsteady loads on the blades are still in the process of development. Theories are available, but complementary experimental information is lacking. Flow induced fluctuating pressures on the blades are believed to be important in the prediction of the higher frequency noise components significant in subjective reaction. These fluctuating pressures may be due to nonuniform in-flow to the propeller disk as a result of atmospheric turbulence, configuration as- symetries, and thrust axis inclination.

Some performance penalties may be expected due to the quieting of general aviation aircraft.

These can take the form of reduced payload because of the additional weight of mufflers and multiblade propellers, reduced cruise speed, reduced range, and in- creased takeoff distance. In several situations, evaluations have been made of per- formance penalties associated with specified amounts of noise reduction (see refs. 76, 79, and 80).

One finding of recent studies is that the propellers currently used on general avia- tion aircraft are not necessarily well matched to the engine and airframe. Thus, an optimized propeller may be able to provide a substantial improvement in noise with a minimal effect on the performance of the aircraft.

S Planned research will be oriented to developing and confirming methods for high- frequency propeller noise prediction. This will involve precision acoustic and aerody- namic measurements in quiet wind tunnel environments, such as the Aircraft Noise Reduction Laboratory, to provide empirical inputs to the theories and to provide overall validations. Some limited in-flight studies are also required for correlation.

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Fisk, W. S.; Hedmann, H. W.; Knott, P. R.; and Sandusky, G. T.: Supersonic Transport Noise Reduction Technology Summary - Phase I. General Electric Co., Report No. FAA-SS-72-43 (Contract FA-SS-67-7), Dec. 1972.

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11. von Glahn, Uwe H.; Gray, Vernon H.; Krejsa, Eugene A.; Lee, Robert; and Minner, Gene L.: Jet Noise.

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12. Gutierrez, Orlando A.; and Stone, James R.: Preliminary Experiments on the Noise Generated by Target-Type Thrust Reverser Models.

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Washington, - Far Field Radiated Aerodynamic The Ultimate Noise Barrier 17. Gibson, John S.: Control Engineering, Noise. Proceedings of the International Conference on Noise 4-6, 1972.

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jointly by the INCE and ASA, sponsored Transport Noise Reduction Tech- J. F. and Doyle, V. L.: Supersonic 18. Brausch, 1972.

I. Vol. 1. FAA-SS-72-42, Dec.

nology Summary - Phase Tech- Transport Nosie Reduction F.; and Doyle, V. L.: Supersonic 19. Brausch, J.

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- Phase I. Vol. 2. FAA-SS-72-42, nology Summary G. T.: Supersonic H. W.; Knott, P. R.; and Sandusky, 20. Fisk, W. S.; Hehman, Summary - Phase I. FAA-SS-72-43.

Noise Reduction Technology Transport SST Technology D. B.; Benefiel, R. L.; and Simcox, C. D.: 21. Lu, H. Y.; Morden, Evaluation of an SST Noise Suppres- Follow-on Program. Phase I - Performance FAA-SS-72-40-1, 1972.

Nozzle System. Vol. 1: Suppressed Mode.

sor SST Technology Follow-On Program. Phase I - Performance 22. Morden, D. B.: of an SST Noise Suppressor Nozzle System. Vol. 2: Unsuppressed Evaluation Modes. FAA-SS-72-40-2, 1972.

Acceleration and Cruise C. P.; Morden, D. B.; and Simcox, C. D.: SST Technology Follow-On 23. Wright, Test Program.

Program. Phase I - A Summary of the SST Jet Noise Suppression 1972.

FAA-SS-72-41, Separate 24. Huff, R. G.; and Groesbeck, D. E.: Splitting Supersonic Nozzle Flow Into Jets By Overexpansion Into a Multilobed Divergent Nozzle.

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Report on Jet Noise Suppression as 25. Swan, W. C.; and Simcox, C. D.: A Status By an Aircraft Manufacturer. Presented at the First International Symposium Seen 1972.

on Air Breathing Engines, Marseilles, France, June 19-20, H.; Gray, Vernon H.; Krejsa, Eugene A.; Lee, Robert; and 26. von Glahn, Uwe Minner, Gene L.: Jet Noise. Aircraft Engine Noise Reduction Conference.

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From Inter- 27. Dosanjh, D. S.; Abdelhamid, A. N.; and Yu, J. C.: Noise Reduction acting Coaxial Supersonic Jet Flows. Basic Aerodynamic Noise Research Con- ference, NASA SP-207, 1969.

28.

Brausch, J. F.: Flight Velcoity Influence on Jet Noise of Conical Ejector, Annular Plug, and Segmented Suppressor Nozzles.

(General Electric Co., Evendale, Ohio.) NASA CR-120961, Aug.

1972.

29. Burley, R.

B.; and Karabinus, R. J.: Flyover and Static Tests to Investigate Ex- ternal Flow Effect on Jet Noise for Non-Suppressor and Suppressor Exhaust Nozzles. NASA TMX-68161, 1973.

30. Carlson, Harry W.; McLean, F. Edward; and Middleton, Wilbur D.: Prediction of Airplane Sonic-Boom Pressure Fields.

Conference on Aircraft Operating Problems.

NASA SP-83, 1965, pp. 235-244.

31. Anon.: Proceedings of the Sonic Boom Symposium sponsored by Acoustical Society of America, St. Louis, Missouri, November 3, 1965.

J. Acous.

Soc. Amer., vol. 30, no. 5, May 1966, pp. S1-S80.

32. Seebass, A. R., ed.: Sonic Boom Research.

NASA SP-147, 1967.

33. Miller, David S.; Morris, Odell A.; and Carlson, Harry W.: Wind-Tunnel Investi- gation of Sonic-Boom Characteristics of Two Simple Wing Models at Mach Numbers From 2.3 to 4.63.

NASA TN D-6201, 1971.

34. Schwartz, Ira R., ed.: Third Conference on Sonic Boom Research.

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35.

Schwartz, Ira R., ed.: Second Conference on Sonic Boom Research. NASA SP-180, 1968.

36. Stanford Research Institute: Sonic Boom Experiments at Edwards Air Force Base.

NSBEO-1-67 (Contract AF 49(638)-1758), NTIS, U.S.

Dep. Com., July 28, 1967 (Available from DDC as AD 655 310).

37. Miller, David S.; and Carlson, Harry W.: A Study of the Application of Heat or Force Fields to the Sonic-Boom-Minimization Problem.

NASA TN D-5582, 1969.

38. Garrick, I. E.; and Maglieri, D. J.: A Summary of Results on Sonic-Boom Pressure-Signature Variations Associated With Atmospheric Conditions.

NASA TN D-4588, 1968.

39. Maglieri, Domenic J.; Hilton, David A.; and McLeod, Norman J.: Experiments on the Effects of Atmospheric Refraction and Airplane Accelerations on Sonic- Boom Ground-Pressure Patterns.

NASA TN D-3520, 1966.

40. Vallee, J.: Mesure de l'Intensite des Bangs Soniques Engendres par un Avion Volant en Palier Accelere Supersonique.

Rapp. d'Etude No. 272, Centre d'Essais en Vol, Annexe d'Istres, Oct. 1967.

des Focalisations de Bangs Soniques Engendrees 41. Vallee, J.: Etude Experimentale ou en Virage d'un Avion Vol Supersonique en Acceleration Rectiligne par le 277, Centre d'Essais IV a l'Altitude de 11, 000 m. Rapp. d'Etude No.

Mirage Vol, Annexe d'Istres, May 1969.

en the Second Sonic Boom 42. Ribner, H. S.; and Hubbard, H. H., eds.: Proceedings of 1972.

Symposium, American Institute of Physics, 43. Randall, D. G.: Methods for Estimating Distributions and Intensities of Sonic 1959.

Bangs. R. & M. No. 3113, Brit. A.R.C., 44. Kane, Edward J.; and Palmer, Thomas Y.: Meteorological Aspects of the Sonic Boom. SRDS Rep. No. RD64-160 (AD 610 463), FAA, Sept. 1964.

45. Haglund, G. T.; and Kane, E. J.: Flight Analysis of Sonic Boom Phenomena Near the Shock Wave Extremity. NASA CR-2167, 1973.

46. STOL Technology Conference, Ames Research Center, October 1972. NASA SP-320, 1973.

47. Rulis, R.: Status of Current Development Activity Related to STOL Propulsion Noise Reduction. NASA TM X-68195.

48. Quiet Turbofan STOL Aircraft for Short-Haul Transporation: Lockheed; Interim Report Phase I Completion. October 1972.

49. Study of Quiet Turbofan STOL Aircraft for Short-Haul Transportation: Douglas Aircraft Company. Phase I Report. October 1972 ( 6 volumes).

50. Kelley, G. S.; and Gerend, R. P.: Propulsion Systems for Commercial STOL Aircraft. AIAA Paper 71-746, June 1971.

51. The Boeing Company: Design Integration and Noise Study for a Large STOL Aug- mentor Wing Transport - Task I Report. Rept. D6-60139, July 1971.

52.

"Design Integration and Noise Studies for Jet STOL Aircraft. " Task V Final Report. NASA CR-114534, 1973.

53. Koenig, D.; Corsiglia, V.; and Morelli, J.: Aerodynamic Characteristics of a Large Scale Model with an Unswept Wing and Augmented Jet Flap. NASA TN D-4610, 1968.

54.

"Design Integration and Noise Studies for Jet STOL Aircraft. " Final Report, May, 1972: Volume I, NASA CR-114283; Volume II, NASA CR-114284; Volume III, NASA CR-114285; Volume IV, NASA CR-114286.

55. Dorsch, R. G.; Krejsa, E. A.; and Olsen, W.

A.: Blown Flap Noise Research.

AIAA Paper No. 71-745, NASA TMX-67850, 1971.

56. Hardin, J. C.; Kreim, W. J.; and Dorsch, R. G.: Dynamic Pressure Measure- ments on a Half-Scale Externally Blown Flap Model. NASA L.W.P. 989, Part E, Sept. 1971.

57. Goodykoontz, J.

H.; Olsen, W. A.; and Dorsch, R. G.: Small-Scale Tests of the Mixer Nozzle Concept for Reducing Blown Flap Noise. NASA TM X-2638, 1972.

58. Reshotko, M.; Olsen, W. A.; and Dorsch, R. G.: Preliminary Noise Tests of the Engine-Over-the-Wing Concept. I. o0-600 Flap Position. NASA TM 3 0 X-68032, 1972.

59. Dorsch, R. G.; and Reshotko, M.: EBF Noise Tests with Engine Under- and Over- the-Wing Configuration.

STOL Technology. NASA SP-320, 1973.

60. Olsen, W. A.; Miles, J. H.; and Dorsch, R. G.: Noise Generated by Impingement of a Jet Upon a Large Flat Board. NASA TN D-7075.

61. Dorsch, R. G.; Reshotko, M.; and Olsen, W. A.: Flap Noise Measurements for STOL Configurations Using External Upper Surface Blowing. NASA TM X-68167.

62. Dorsch, R. G.; Kreim, W. J.; and Olsen, W. A.: Externally Blown Flap Noise.

AIAA Paper No. 72-129, Jan. 1972.

63.

Olsen, W. A.; Dorsch, R. G.; and Miles, J. H.: Noise Produced by a Small-Scale Externally Blown Flap. NASA TN D-6636, 1972.

64. Goodykoontz, J. H.; Dorsch, R. G.; and Groesbeck, D.

E.: Mixer Nozzle Exter- nally Blown Flap Noise Tests. NASA TM X-68021, 1972.

65. Dorsch, R. G.; Lasagna, P. L.; Maglieri, D. L.; and Olsen, W. A.: Flap Noise.

NASA SP-311, 1972.

66. Helms, H. E.: Quiet Clean STOL Experimental Engine Study Program. Task I - Parametric Propulsion System Studies. Sept. 1972.

67.

Helms, H. E.: Quiet Clean STOL Experimental Engine Study Program. Task II - Preliminary Design Studies. To be published in 1973.

68. QCSEE Study: General Electric Company, Task I. To be published in 1973.

69. QCSEE Study: General Electric Company, Task II. To be published in 1973.

70. Sweet, H. S.: Design Studies of Low-Noise Propulsive Lift Airplanes.

SAE Air Transportation Meeting, Miami, Florida, April 24-27, 1972 (to be published).

71. Davis, D. D., Jr.: Acoustical Filters and Mufflers, ch.

21 of Handbook of Noise Control.

C. M. Harris, ed., McGraw-Hill Book Company, Inc., New York, 1957.

Chamber Muffler Design with Method for Expansion 72. Parrott, T. L.: An Improved TN.

Proposed NASA Operational Muffler.

Application to an Charles E.: Trends in Noise Harvey H.; and Feiler, David; Hubbard, 73. Chestnutt, Aircraft Plants. Presented at NASA for Aircraft Gas Turbine Power Control Center, May 4-6, Safety and Operating Problem Conference, Langley Research SP-270.

1971. NASA L.: A Review of L.; and Runyan, Harry Harvey H.; Lansing, Donald 74. Hubbard, vol. 19, no. 3, Rotating Blade Noise Technology. J. Sound and Vibration, 1971, pp. 227-249.

Dec.

David A.: Results of a Noise Robert J.; and Hilton, Henderson, Herbert R.; Pegg, 75.

NASA TN Modified OH-6A Helicopter.

Program on a Standard and Measurement 1973.

D-7216, Reduce of a Light Airplane Modified to Sound-Level Measurements 76. Vogeley, A. W.: NACA TN 1647, 1948.

Noise Reacting the Ground.

and Taylor, C. F.: Experiments F. S.; Roberts, J. P.; 77. Beranek, L. L.; Elwell, TN 2079, May 1950.

of Light Airplanes. NACA in External Noise Reduction in External Noise Reduction of Beranek, L. L.: Experiments 78. Roberts, J. P.; and July 1952.

Airplane. NACA TN 2727, a Small Pusher-Type Amphibian Phenomena Relating D. J.: An Investigation of Some Hubbard, H. H.; and Maglieri, 79.

1958.

of Airplanes. NACA TN 4337, to Aural Detection Applications of Quiet Aircraft D.; and Roberts, G. F.: Commercial 80. Griffith, E.

at SAE National Business Aircraft Meeting, Wichita, Technology. Presented Kansas, March 1972.

Its of Research on Propeller Noise and A.; and Hubbard, H. H.: Status 81. Regier, A.

3, May 1958, pp. 395-404.

Acous. Soc. Amer., vol. 25, no.

Reduction. J.

of Rotating Blade D. L.; and Runyan, H. L.: A Review 82. Hubbard, H. H.; Lansing, Noise Technology. J. Sound and Vibration, vol. 19, no. 3, Dec. 1971, pp. 227-249.

BIBLIOGRAPHY D3-9039.

and Airplane Integration Definition Boeing: 707/JT3D-9 Refan Nacelle November 10, 1972.

Integration Definition D3-9039-1, Boeing: 707/JT3D-9 Refan Nacelle and Airplane Submittal. January 15, 1973.

Second Pratt & Whitney: Preliminary Engine Definition and Characteristics of the JT3D Quiet Engine. September 14, 1972.

Pratt & Whitney: Preliminary Engine Definition and Characteristics of the JT8D Quiet Engine.

October 14, 1972.

Pratt & Whitney: Final Phase I Engine Definition and Characteristics of the JT3D-9 Engine. December 14, 1972.

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Report MDC J5731A, January 8, 1973.

McDonnell Douglas: DC-8 Series 63 Engine and Nacelle/Airframe Integration Definition.

Report MDC J5732, November 10, 1972.

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Report MDC J5732A, January 8, 1973.

McDonnell Douglas: Preliminary Retrofit and Economic Analysis, Volume I Economic Analysis.

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Report MDC J5734, January 8, 1973.

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page

Preceding

III -OPERATING

PROCEDURES FOR

AIRCRAFT NOISE REDUCTION

Operational procedures can be used effectively for noise control in both landing- approach and the takeoff-climbout phases of the mission. The interrelated factors of aircraft altitude, engine throttle setting, flap angle setting, and aircraft speed are significant.

NASA, in cooperation with FAA and the airlines, has been involved in developing and evaluating operational procedures for noise reduction for a number of years, both for takeoff-climbout and landing-approach situations. The takeoff-climbout studies (refs. 1 to 4) have been helpful in evaluating the noise reduction potential for various flap angle and engine throttle schedules for a number of aircraft. These data have also been use- ful as a guide in defining the optimum procedures for particular operations.

A main finding of these takeoff-climbout studies is that the optimum conditions for noise alleviation depend on the configuration details (particularly, type of engine) and operating characteristics of the aircraft and thus will probably be different for each new aircraft.

The landing-approach studies on the other hand have indicated potentially larger noise reductions, and they are not so configuration oriented.

Three noise reduc- tion techniques that have been proposed are the two-segment approach, the energy management or decelerating approach, and the curved ground track approach.

The two-segment approach concept is illustrated in figure

r-1. The upper

profile represents the two-segment approach, and the lower profile is a standard instrument landing approach.

Using the two-segment approach, the aircraft approaches on a steeper TWO- SEGMENT APPROACH NORMAL APPROACH RUNWAY Figure III-1. - Two-segment approach concept path in than normal glide slope and then makes a transition to the standard approach By keeping the aircraft higher above the ground and time to stabilize before -landing.

of the steeper angle, the two-segment approach reducing the engine power because lessens the community noise near airports.

initiates the ap- In the energy management or decelerating approach, the aircraft to landing speed at at a relatively high airspeed and then slowly decelerates proach of the reduced power, the noise under the approach greatly reduced power. Because has the potential path is reduced. The decelerating approach is attractive because it all the way to the threshold. This technique might be of providing some noise relief best feature of each.

combined with the two-segment approach in order to use the noise sensitive areas by approaches on a The third procedure is based on avoiding today. With curved ground track. This technique is being used under visual conditions technique can be the advent of area navigation and the microwave landing system, this and combined with the two-segment approach.

extended to instrument flight conditions Although these noise abatement flight procedures are well within the performance capability of current day jet transports, they impose new requirements on the pilot on the guidance and navigation system, on the duties and workload, on the pilot displays, aircraft control system, on Air Traffic Control (ATC) flow of aircraft to high density runways and on parallel runway operations, and possibly different wake turbulence ef- to develop suitable avionics for noise fects. A substantial effort is therefore required abatement procedures and to obtain sufficient experience so that they are accepted for routine operations.

of this report, the NASA program directed towards developing For the purpose operational procedures for noise abatement is divided into two parts. The first part is that will allow aircraft to aimed at developing operational avionics and flight procedures during routine make two-segment approaches under instrument flight conditions scheduled operation. This part of the program is currently under way, and significant the feasibility of progress has been made. The second part is aimed at determining other techniques for noise abatement such as the decelerating approach or curved ground how to track approach. The second part of the program also addresses the problem of new navigational aids such as the microwave landing system. Work related best utilize initiated.

to the second part of the program has not yet been PROGRAM HISTORY studies to obtain a preliminary determi- The FAA and NASA have conducted several nation of the feasibility of using modified operating procedures to reduce the noise perceived by the airport community. Both agencies have determined that significant noise reduction can be achieved by using the two-segment approach.

NASA has been primarily concerned with the evaluation of pilot displays that would be required to make noise abating two-segment approaches (refs.

5 to 8). The FAA has been primarily con- cerned with developing the necessary guidance systems (refs. 9 and 10).

In these studies, experimental equipment was evaluated to assess concept feasibility.

NASA and American Airlines recently completed a program to incorporate the results of the previous studies into operational equipment.

The goal of the program was to assess the operational feasibility of the two-segment approach as a method of reducing airport community noise (ref. 11).

For these tests, an area navigation system was used to compute the upper segment, and the instrument landing system (ILS) glide slope was used for the lower segment.

The localizer was used throughout the approach.

A key feature of the program with American Airlines was the provision of a continuous vertical steering command on the flight director. This was required to insure that transitions from level flight to the upper segment could be made without overshoots and those from the upper to the lower segment could be made without going below the normal ILS.

The additional power needed to correct for going below the ILS is particularly objectionable because it creates higher perceived noise on the ground in the region of the transition.

This effect is illustrated in figure 1T-2.

The tests with American Airlines were conducted during a 30-day period in the summer of 1971 at the Stockton, California, Metropolitan Airport.

Stockton Metropolitan Airport was selected for these tests because of the low traffic density and good visibility 244- 800 APPROACH ANGLE, / 183- 600- DEC _ - - ALTITUDE, 122- ALTITUDE, 00 M F -2. 65- / 61- 200 -, / CHANGE IN __AT NOISE LEVEL 0 PNdB 101 -. 5 .5 1.0 1.5 2O DISTANCE FROM RUNWAY THRESHOLD, N. MI.

I I I I I I -. 9 0 .9 1.8 2.8 3.7 DISTANCE FROM RUNWAY THRESHOLD, KM Figure III-2. - Effects of inadequate guidances.

prevalent during the test period.

The program demonstrated that two-segment ap- proaches might be operationally feasible and warranted a much more extensive and thorough evaluation under actual operational conditions.

The results of the program with American Airlines were presented to the NASA Research and Technology Advisory Committee on Aeronautical Operating Systems and to the Ad Hoc Panel on Noise Abatement by Operational Procedures.

These advisory committees are composed of individuals representing the airlines, airframe manufac- turers, avionics suppliers, the Air Transport Association (ATA), the Air Line Pilots Association (ALPA), FAA, and DOT.

The committees agreed that the two-segment approach appeared operationally feasible and warranted additional evaluation.

They recommended that further flight evaluations be conducted under representative opera- tional conditions in two aircraft types: A Boeing 727 aircraft, because these aircraft account for the largest number of arrivals and departures and are owned by more air carriers than any other aircraft, and a long-range aircraft such as the DC-8 or Boeing 707 because these aircraft differ significantly from the Boeing 727 and have a larger noise footprint.

The panel also recommended that the results of these two flight pro- grams be extrapolated through analysis and simulation to determine the applicability of the two-segment approach to the other aircraft in today's fleet.

TWO-SEGMENT APPROACH The first part of this program consists of several steps. The first two steps are being conducted with United Air Lines and call for separate flight evaluations using a Boeing 727-200 and a McDonnell-Douglas DC-8-61, each equipped with different avionics for providing vertical guidance during the approach.

The Boeing 727 will be equipped with a special purpose glide slope computer, and the DC-8 will be equipped with an area navigation system.

Both systems will be designed and built by the Collins Radio Company under contract to NASA.

The glide slope computer system is being evaluated as an inexpensive retrofit for aircraft not equipped with area navigation equipment.

The area navigation system is being evaluated to determine the operational feasibility of modifying the existing airborne area navigation equipment to provide the two-segment capability. If the aircraft has an installed area navigation system, this concept appears to be the least expensive way to add the two-segment approach capability.

-Another step in this part of the program involves the extension of the flight results to the other air- craft in today's fleet.

STEP A: DEVELOPMENT AND FLIGHT EVALUATION OF A SPECIAL PURPOSE GLIDE SLOPE COMPUTER IN A BOEING 727-200 AIRCRAFT NASA Ames Research Center began work on this program with United Air Lines and the Collins Radio Company in July 1972. The program objectives are to develop an inexpensive avionics retrofit kit that will make an aircraft capable of a two-segment approach and to evaluate the two-segment approach in a Boeing 727-200 aircraft during regular scheduled service.

The program includes avionics design and fabrication; a simulation study aimed at developing a procedure and profile that is safe under adverse conditions; an engineering flight evaluation devoted to equipment checkout, certification, and verification of the approach profile established during the simulation study; a 1-month series of off-line flight evaluations; and a 6 -month evaluation in revenue service.

The avionics design and fabrication, the simulation study, the engineering flight evaluation, and the off-line pilot evaluation have been completed.

The results of these phases have not been completely reviewed and analyzed, but preliminary indications are that the avionics and two-segment approach are operationally feasible in the Boeing 727 and acceptable to the airline community.

In the simulation study the task was to make the concept into a practical, operational reality since the basic concept of the two-segment approach had been established by previous studies and research projects.

In the design of the two-segment procedures, the basic profile was divided into eight parts as illustrated in figure I1-3.

The effect of UPPER 1 CAPTURE POINT-, I L. LUPPER" TRANSITION UPPER SEGMENT, UPPER rLOWER CAPTURE ANGLE-' SEGMENT POINT INTERSECT ALTITUDE LOWER LOWER rGLIDE SCOPE INTERSECT ALTITUDEN \DME RUNWAY- Figure III-3.

- Noise abatement approach profile simulation variables.

parts are was examined. Comments regarding these eight each part on the approach contained below: intercept altitude - The system must function such that this part (1) Upper segment to 6000 feet (ft) altitude flight level (AFL) (and even higher is desirable). Also, can vary climbing or descending.

it must not be fixed but either to vary from 1500 ft AFL down to Lower intersect altitude - This part was made (2) but it was height. A practical operational range would be smaller, runway threshold over this range.

felt that its influence on the approach should be tried made to vary from 40 to 70, although 80 (3) Upper segment angle - This part was previous information about these descent and 100 were added to check the validity of angles.

- This part was expanded from the nominal glide slope range of (4) Glide slope ° .

a bias allowing the 2.50 to 3.00 to 3. 5 The system was designed so as to provide to hold the additional angle increment over the standard ILS glide pilot to have guidance slope.

to the pilots capture point - This part was considered very important (5) Upper so designed to compensate for varying acceptance and passenger comfort. It was rates to the upper segment angle.

closure was also considered important to safety, pilots (6) Lower capture point - This part acceptance, and passenger comfort. It was designed to compensate for varying closure rates to the glide slope.

to passenger comfort, was designed to (7) Upper transition - This part, important allow wide variations that enable the pilot to get to the upper segment without additional the passengers.

constraints or disturbances to - This part was considered the key to pilot acceptance and (8) Lower transition was designed so that the pilot could make this transition using a normal instrument close check and normal flight technique, and not feel that he was performing an unusual maneuver that would require him to restabilize the aircraft at its completion.

the external variables that the pilots might encounter were The effects of some of examined in the simulation. A summary of some of these are listed here: Turbulence - The two-segment approach during simulation was not adversely (1) was flyable on affected by turbulence. Any turbulence level flyable on the standard ILS the two-segment approach. In the airplane the two-segment approach required less was significant turbulence.

effort than the standard ILS when there (2) Icing - With engine and wing anti-icing on and temperatures -70 C or above, the low pressure turbine rpm is about the minimum of 55 percent. In these conditions a using 400 flaps. But if the icing is such that tail wind of about 15 knots can be offset by 70 percent N is required for anti-icing, or the tail winds are in excess of 15 knots, then the approach, as constituted, could not be flown. These conditions exist less than 1 percent of the time.

(3) Winds - Tail winds in excess of 30 knots present a problem of airspeed stabili- zation and throttle position.

Less than 30 knots are maneuverable.

Cross wind effect is the same as the standard ILS. Wind shear effect is very similar also, except that the upper segment can be followed easier than the glide slope when troublesome wind shear is present.

(4) Visibility - No noticeable difference between the two-segment approach and the standard ILS was detected.

(5) Lighting - The two-segment approach profile permits a better view of the terminal area under all lighting conditions than does the standard ILS, yet the descent angle is not so steep as to give the pilot the impression of his descending into a hole at night.

(6) Airports - The relationship of the two-segment approach and the standard ILS is very similar at Los Angeles, San Francisco, and Stockton.

(7) Navaid failures - No difference, except that the colocated distance measuring equipment (DME) adds in one more system that must be in operation for the two-segment computer to function.

The two-segment approach that resulted from the simulation evaluation was used in the engineering flight evaluation.

The upper intersect altitude was designed to go as high as 6000 ft AFL.

The altitude was tested and found successful up to 14 000 ft (mean sea level). The upper and lower capture points occurred as designed and were very satisfactory.

The upper segment angle was selected to be 5.20 to 7.00.

The lower value was found to have good noise improvement when associated with low-lower inter- sect altitudes.

It also allowed the Boeing 727 to use full anti-ice capability when 400 flaps were used.

The upper value was determined to be the greatest angle expected at any time during any two-segment approach with a Boeing 727.

The Supplemental Type Certificate (STC) demonstrations were made at this angle.

The glide slope angle will be the same that the ILS has for the airport concerned.

The values 2.

50 to 3. 50 covers all ILS glide slope angles that would be of concern.

The system is capable of flying high on the glide slope with a fixed bias. This was flown during the engineering flight evaluation and was found to have merit, but it will not be used during the on-line flight evaluation.

The lower intersect altitude range was 400 to 800 ft AFL.

The nominal value determined by flight evaluation was about 700 ft.

The ground noise measurements were made at the high and low values of this range.

The two-segment approach profile, resulting from the flight evaluation, was used for the off-line pilot's evaluation and is basically the same as will be used for the on-line pilot's evaluation.

The Stockton, California, profile is shown in figure I-4. The San Francisco and Los Angeles profiles are very similar.

The angle of the standard ILS is different, and THE GLIDE SLOPE WILL NOT CAPTURE WHEN THE AI RCRAFT IS BELOW A UPPER SEGMENT DOT ABOVE GLIDE SLOPE WITH NO IF AIRCRAFT WITHIN ONE-HALF B DISENGAGEMENT GLIDE SLOPE CAPTURE NO GLIDE SLOPE CAPTURE AIRCRAFT WITHIN 2.2N.MI. DME WITH ( DISENGAGEMENT IF APPROACH ALTITUDE 15 000 FT AFL MAXIMUM ALTITUDE 3100 FT MSL * MINIMUM APPROACH UPPER SLOPE, UPPER CAPTURE POINT; r-.) LOWER CAPTURE POINT; 1050 FT MSL, 130 KNOTS IAS 6. 4DME, 160 KNOTS I AS;.

DAY CALM WINDS STANDARD CALM WINDS ., STANDARD DAY GIESLOPE, Z

GLIDE

ON GLIDE SLOPE WITHIN

1/4 DOT 550 FT MSL

* pLOPE STOCKTON CALIFORNIA TDZ 29 FT MSL used at Stockton, California, resulting from flight evaluation.

Figure III-4. - Two-segment approach profile The and the lower capture point.

of the lower intersect altitude this results in a shift airspeed, is about 100 ft flown at the lowest shift with the lowest angle glide slope, high as captured and flown very satisfactory, as lower. The upper segment can be not necessary for guidance on the Localizer capture or alignment is 15 000 ft AFL.

upper segment.

of the profile to increase the flight Safety factors were designed into some areas set, the DME, or the airport safety margins for the approach. In the event the baro presented prematurely. To elevation panel set malfunctions, the upper segment could be cause a descent below the standard glide slope, prevent a guided approach that would glide slope.

from capturing when the aircraft is below the the upper segment is prevented and gets to within one-half dot deflection above If the aircraft is flying the upper segment bias out of view.

auto pilot will disengage and the flight director bars the glide slope, the the guidance that would take the aircraft below This prevents the system from providing is presented late, it would be possible to descend so glide slope. If the upper segment case the system will slope would be reached very low or not at all. In that that the glide 2.2 nautical miles DME and the glide slope is not disengage if the aircraft reaches captured.

If the pilot can get upper and lower transitions were a key to pilot acceptance.

The without any significant change in his flight technique, into and out of the upper segment sound.

he should accept the two-segment concept as operationally If the aircraft is approaching The upper transition starts at the upper capture point.

a low or is climbing, the capture point occurs early. If the aircraft is at at a high speed late. In either case, the aircraft is pitched speed or is descending, the capture occurs nose down slowly and smoothly, such that the upper segment is reached in 500 to 800 ft below the initial altitude at capture.

The lower transition is a smooth, easy pitch change that starts at the lower capture point.

The lower capture point will adjust according to the speed at which the aircraft is closing on the glide slope.

At high speeds the capture occurs earlier and provides a more gradual pitch change than at low speeds.

The result is that the transition seems similar to both pilot and passengers.

Passengers do not detect the lower transition.

The point at which the glide slope is reached does not shift to any great extent.

The upper segment tracking with its transitions was determined to be very satisfac- tory.

It required no additional pilot skills for routine operation of the Boeing 727-200 aircraft.

The off-line evaluation consisted of a two-phase program to thoroughly familiarize the guest pilot with the two-segment approach, thereby enabling him to evaluate the approach in detail.

Phase I was the viewing of an audio-visual package followed by a crew briefing and a 1-hour and 30-minute simulator flight.

The simulator involved a syllabus of 11 approaches intermixing the standard ILS with the two-segment ILS under varying weather conditions and operational techniques.

Phase II consisted of an air- craft period during which an eight approach syllabus was flown, which again compared the standard ILS with the two-segment ILS in a real world environment.

The expected 90- and 95-effective perceived noise decibels (EPNdB) contours for a Boeing 727-200 aircraft using this two-segment approach procedures are compared in figures III-5 and III-6 with the contours expected as a result of using a standard instru- ment landing approach.

The 90-EPNdB impacted area is reduced during the two-segment approach by 3.7 square miles (67 percent reduction).

The 95-EPNdB impacted area is reduced by 1. 1 square miles (48 percent reduction).

By increasing the upper intersect altitude, there can be a significant improvement in ground noise outside the outer marker. Altitude of up to 6000 ft AFL can produce noise improvement over large areas in approaching the airport.

The aircraft safety is enhanced by staying high in the heavy traffic area, which reduces exposure to many low flying aircraft.

It was noticed that the approach with a 60 upper segment could accom- modate up to 190 knots (indicated air speed) at 3000 ft AFL to the point of upper segment capture.

This speed can be increased as altitude increases up to 250 knots at 6000 ft AFL or higher.

The result is lower power setting at higher altitudes and less time at high power settings.

This could produce a side benefit of lower fuel consumption of each approach.

The avionics system being evaluated by United Air Lines retains the coupled flight director feature used in the American Airlines program and adds the autopilot coupling so that the pilot can make a two-segment landing with all the aids available for standard approaches.

,-STANDARD INSTRUMENT LANDING MI.

2.1 SQ.

/APPROACH, LANDING INSTRUMENT ,,-STANDARD SAPPROACH, 5.5 SQ.

MI. -TWO-SEGMENT APPROACH, . ,,TWO SEGMENT APPROACH, 1.0 S. MI.

.5 M'../8 OSQ.5 i i i I I I I I I I I I -10 -8 -6 -4 -2 0 2 -6 -4 -2 0 2 DISTANCE, N.MI. DISTANCE, N.MI.

Figure III-5. - 90-EPNdB approach contours for Boeing 727. Figure III-6. - 95-EPNdB approach contours for Boeing 727.

of the two-segment system stressed adherence to standard United's implementation to such an extent that one-switch operation and an airport elevation input procedures standard are the only features that distinguish the two-segment procedure from United's ILS procedure.

The special purpose glide slope computer developed by Collins uses a signal from a DME transmitter colocated with the ILS glide slope and barometric corrected pressure aircraft on the upper segment and uses the ILS glide slope devia- altitude to position the tion to position the aircraft on the lower segment. The two-segment computer also uses altitude rate information from the Central Air Data Computer (CADC) for vertical to drive an autothrottle.

path damping and airspeed from the CADC DME transmitters, colocated with the ILS glide slope, are not standard equipment in an instrument landing system. However, the FAA currently plans to add these facilities at a rate of five in FY 75, 50 in FY 76, 30 in FY 77 and 40 in FY 78. The necessary colocated facilities are available at the airports being used in the program.

Although it is very difficult to estimate the cost of retrofitting United Air Lines fleet of Boeing 727's with this system, it is thought that the cost will be approximately $31 400, for a dual installation. The $31 400 assumes $26 600 for equipment, $4000 for installation, and $800 for flight check. Out-of-service and training costs are not included. It is assumed that installation could occur when the aircraft are out of service for other reasons and that training could be incorporated into the normal training and review curriculum.

For several reasons, the present program is providing a much broader basis for evaluating the feasibility of the two-segment approach than in previous programs. First, the avionics have been designed, built, and environmentally tested to FAA Technical Standard Order specifications. The system performs internal selfchecks and, in the event of a failure, provides the pilot with a warning similar to warnings provided in the event of a failure during an ILS approach. Second, the procedure and system have been tested both in the simulator and in flight under a wide variety of operational condi- tions.

Approaches have been made under instrument flight conditions; in the presence of tail winds, wind shears, and turbulence; at dusk and at night; and at several airports including Los Angeles and San Francisco.

Third, over 50 pilots have participated in the off-line pilot evaluation: 15 line pilots representing ALPA and APA, 19 management pilots from the different airlines, 11 FAA pilots, five engineering test pilots, and one USAF pilot. Finally, a broader spectrum of line pilot reactions will be obtained as a result of the in-scheduled service evaluation, which begins in late April 1973 and lasts through October 1973. This will be the first time a two-segment guided approach sys- tem has been placed into routine line service.

During this period it is expected that over 96 crews will evaluate the system and that over 500 two-segment approaches will be made.

STEP B: DEVELOPMENT AND FLIGHT EVALUATION OF TWO-SEGMENT AVIONICS USING THREE-DIMENSIONAL AREA NAVIGATION FOR GUIDANCE IN A DC-8-61 United Air Lines and the Collins Radio Company initiated work, under contract with NASA, on this program in December 1972.

The program objectives are to determine the operational feasibility of modifying a three-dimensional area navigation system to provide the two-segment approach capability and to evaluate the two-segment approach in a DC-8-61 aircraft in regular scheduled service.

The program contains the same basic phases as the Boeing 727 evaluation covered in STEP A. However, the avionic concept and aircraft characteristics are substantially different.

In this step an existing area navigation system will be modified to include the two- segment capability.

An inherent advantage of this concept is that, if the aircraft is equipped with an area navigation system, a modification to the system represents an inexpensive way of incorporating the two-segment approach capability.

A second advantage is that the system can be used to make precision approaches to ILS equipped runways without requiring a colocated DME transmitter facility. The system can also be used to make nonprecision noise abating approaches into non-ILS equipped runways.

The Boeing 727 aircraft used in STEP A is particularly well suited for the two- segment approach. It has relatively high drag in the landing configuration and requires positive thrust component to come down the 60 glide slope at reference velocity.

It is also equipped with relatively new and complete avionic systems so that the two-segment guidance interface with the autopilot and flight director is straight forward.

On the other hand, the McDonnell-Douglas DC-8 has relatively little drag in the landing configuration and requires near idle thrust to come down a 60 glide slope at has an auto- to the low drag characteristics, the DC-8 reference velocity. In addition preliminary flight tests indicate the Boeing 727 autopilot. Even though pilot older than two-segment guidance command, the interface that the DC-8 autopilot can follow the may require more extensive two-segment guidance system and the autopilot between the it is the opinion on the Boeing 727. For these reasons, modifications than are required evaluation must be con- FAA, and the pilots that the two-segment of the airlines, the approach order to establish the envelope of acceptable two-segment ducted in the DC-8 in of commercial aircraft.

profiles for the fleet approach, the DC-8 is more difficult to adapt to the two-segment Although the DC-8-61 The 90- and 95-EPNdB contours for a expected noise benefits are significant.

with a 690-ft intercept altitude are com- aircraft during a 60/30 two-segment approach for a standard instrument in figures 111-7 and 111-8 with noise contours estimated pared area is reduced by 6.3 square miles (54 landing approach. The 90-EPNdB impacted miles impacted area is reduced by 3. 3 square percent reduction), and the 95-EPNdB (50 percent reduction).

already equipped with area navigation Cost estimates to provide a fleet of aircraft out in detail. However, the with the two-segment capability have not yet been worked glide cost will be substantially less than required to retrofit with the special purpose An estimate of this cost is $ 9000, which includes equipment slope computer system.

It charges. Out-of-service costs and training costs are not included.

and installation could occur when the aircraft are out of service for other is assumed that installation be incorporated into the normal training and review reasons and that training could area navigation curriculum. If the two-segment capability is provided as a part of the quite small.

package prior to installation, it appears that the added cost could become rSTANDARD INSTRUMENT LANDING APPROACH, 11.6 SQ. MI.

,-TWO-SEGMENT APPROACH, 5.3 SQ. MI.

RUNWAY tI I I I I -12 -10 -8 -6 -4 -2 0 2 DISTANCE, N.MI.

Figure 111-7. - 90-EPNdB approach contours for DC-8.

rSTANDARD INSTRUMENT LANDING / APPROACH, 6.6 SQ. MI.

r TWO-SEGMENT APPROACH, MI.

3.288SQ.

I I I I I I -12 -10 -8 -6 -4 -2 0 DISTANCE, N.MI.

Figure III-8. - 95-EPNdB approach contours for DC-& STEP C: STUDY TO DETERMINE THE APPLICABILITY OF THE TWO-SEGMENT APPROACH TO ADDITIONAL JET TRANSPORTS The preceding steps are aimed at determining the operational feasibility of the two- segment approach for only two aircraft types.

The purpose of this step is to extrapolate the results of these flight programs to cover the McDonnell-Douglas DC-9 and DC-10 and the Boeing 707, 737, and 747 jet transports by an analytical and simulation program.

Contracts will be awarded to Boeing and McDonnell-Douglas Aircraft companies in FY 73 to make a preliminary determination of the approach profiles that would achieve maximum noise abatement while maintaining adequate safety margin and pilot acceptance for their different aircraft. These feasibility studies will not include flight simulations.

Contracts will then be awarded to an airline contractor (or contractors) in FY 74 to conduct a simulation study wherein the operational feasibility of making two-segment approaches in these aircraft will be examined in detail.

These studies will look at the effect of extreme wind shear, pilot abuses, and system failures on the safety of the procedure.

STEP D: STUDY TO DETERMINE THE SUITABILITY OF THREE-DIMENSIONAL AREA NAVIGATION TO PROVIDE VERTICAL GUIDANCE An analytical study will be conducted to determine the requirements on the location of the ground navigational aids used as inputs to the airborne navigation equipment in order to provide sufficient accuracy for two-segment guidance.

The study will also define procedures that can be used to flight check the adequacy of existing ground navi- gational aids for establishing the upper segment guidance at individual airports.

with their the FAA in conjunction will be conducted by is expected that this study It area navigation requirements.

aimed at defining existing program TWO-SEGMENT APPROACH THE IMPACT OF THE E: STUDY TO DETERMINE STEP ON ATC with aircraft making have to mix Aircraft making two-segment approaches will approaches for dif- it appears that two-segment ILS approaches. In addition, standard A study will be upper segment glide slopes.

types will require different ferent aircraft approach profiles in on ATC of intermixing different to determine the impact conducted the FAA.

study will be conducted by It is expected that this the terminal area.

FOR NOISE ABATEMENT OTHER TECHNIQUES USING A MICROWAVE OF NOISE ABATEMENT APPROACHES FLIGHT TEST LANDING SYSTEM will have developed 73 considerable expertise and understanding By the end of the FY when flying of the noise abatement operational procedures with respect to the usefulness NAVAIDS, that is, ILS, DME, and the landing approach pattern using the conventional programs FY 73 program and the anticipated follow-on VORTAC. It is hoped that the to carry noise abatement procedures using for FY 74 will provide sufficient momentum in the airlines. Beyond 1974, however, the conventional ground NAVAIDS into practice system, being developed under arises as to the impact of the microwave landing question no real prob- the noise abatement flight procedures. In this respect, FAA contract, on using the microwave landing are anticipated in flying noise abatement procedures lems experience, that is almost inevitable, based on past flight test system. However, it will surface.

certain unanticipated problems wherein noise abatement approaches are Therefore, a flight test program is planned to take advantage of the full using a microwave landing system in an attempt flown could influence the microwave capability of this system and to expose problems that should provide results soon enough system design. Tests conducted in FY 74 landing and development of the microwave system.

to influence the preliminary design how to best use the unique The basic objectives of this program are to determine and to determine if microwave landing system for noise abatement capabilities of the and operational problems associated with there are any navigation, guidance, control, this type of system.

FLIGHT EVALUATION OF CURVED APPROACHES FOR NOISE ABATEMENT Area navigation potentially provides the capability of flying the aircraft along curved approach paths in order to avoid noise sensitive areas. A simulation and flight program is planned, for FY 74 or FY 75, to determine the operational feasibility of using this technique in conjunction with the two-segment approach.

The program will be largely conducted in-house and will include analysis, simulation, and flight test.

A brief des- cription of the effort planned in these phases follows: In this phase, the necessary steering signals will be defined and presentation to the pilot will be evaluated.

Pilot workload and ability to fly these approaches will offer the greatest obstacle. A principal purpose of the simulation will be to determine the amount of automation required to keep the workload at a level comparable with that required during a standard instrument approach.

The effects of winds, wind shears, and pilot abuses will be evaluated.

Flight tests will be conducted using the NASA research Boeing 737 aircraft in the Terminal Configured Vehicle and Avionics Program at the NASA Langley Research Center.

NOISE ABATEMENT USING DECELERATING APPROACHES Two modifications to the standard approach procedure can be proposed for reducing the noise. One consists of flying a steeper-than-standard approach path (i.

e., two- segment approach), which increases the aircraft's altitude over the noise sensitive area and reduces the thrust used in the approach. The other is to make a decelerating approach on a standard glide slope with the engines at idle power.

In this method, the aircraft begins the approach at relatively high airspeed and then slowly decelerates to the landing speed, using the kinetic energy as a power source to overcome the drag forces.

A third method is also possible by combining the two.

If we assume that the approach is flown along the standard ILS glide slope, then, in principle, the decelerating approach can be started at any point on the ILS beam.

The single most important variable in a decelerating approach is the airspeed of the aircraft at the starting point.

This airspeed must be chosen such that the aircraft can fly safely from the outer marker to a desired point with all engines operating at minimum permis- sible thrust, with arrival at.the specified point with full flaps, and with the desired land- ing speed.

Assuming the aircraft arrives at the starting point with the proper airspeed, it begins its gliding and decelerating flight along the ILS beam while either the pilot or an automatic landing system maintains the aircraft's flight along the beam.

As the air- craft is slowly decelerating, the flaps are extended according to a computed schedule.

The novelty of the proposed technique lies in the use of flap angle modulation rather than the more commonly encountered thrust modulation as a method of deceleration control.

If the proper airspeed was selected at the starting point and if the flaps are extended at the proper rate, the landing speed and the full-flap configuration will be reached close to the interception of the glide path with the runway or at any other point along the glide path designated at the terminal point of the deceleration. Since this procedure allows thrust to be maintained at the lowest possible value throughout the approach, engine noise is kept to a minimum. There are safety questions related to this approach because of the time required to spool up the engines if a go-around is required.

ANALYSIS In this phase, the principle objectives are to make a preliminary evaluation of the profile to be flown; that is, whether the decelerating approach should be flown along the o o standard ILS glide slope or along a two-segment glide slope; perhaps along a 3* to 4 glide slope and then about a mile from the runway threshold transition to the normal ILS glide slope. In this phase, the optimum speed profile, flap extension schedule, transi- tion point, flight director requirements for aided manual guidance, guidance laws and interfaces with autopilot and autothrottle for automatic approach, as well as the naviga- tion requirements must be determined.

Piloted Simulation Pilot workload and ability to fly these trajectories will offer the greatest obstacle.

Considerable automation will be required to keep workload from increasing beyond that of standard approaches. A principal purpose of the simulation will be to determine the minimum level of automation needed to keep the workload reasonable.

The simulation program will also evaluate cockpit displays, check out flight director guidance laws and automatic guidance, determine missed approach procedures, study the effect of gusts and wind shears, and define pilot procedures for the manual approach.

Flight Test It is planned that the flight test program will be conducted using NASA Boeing 737 aircraft in the Terminal-Configured Vehicle and Avionics Program.

The main objective of the flight test phase will be to refine the operation of the "decelerating approach" system, further develop the operational procedures, and assess system performance in the actual flight environment.

The final objective of course, is to reduce this experi- mental approach technique to practice.

AERODYNAMIC NOISE Recent computations and measurements have suggested that there may be an aero- dynamic noise floor in the approach and landing configuration of large jets about 10 PNdB below the FAR Part 36 noise level.

Operational procedures such as the two-segment and curved ground track approaches, which increase the separation of the observer and the aircraft, are effective at reducing the impact of aerodynamic as well as engine noise.

The aerodynamic noise varies as a high power of the flight speed. Therefore, the de- celerating approach, which approaches at higher speed, would have a higher aerody- namic noise floor.

In order to obtain better data on the aerodynamic noise floor and understand the re- lationship between aerodynamic noise and engine noise and the different types of noise abatement approaches (the steep glide slope, the two-segment approach, the curved ground track, and decelerating approach) NASA Ames is planning a flight test program with the NASA CV-990, four-engine jet aircraft and possibly other aircraft.

REFERENCES 1. Hilton, D. A.; Copeland, W. L.; and Dibble, A. C., Jr.: Measurements of Noise Produced by a BAC-111-400 Series Turbofan Transport Airplane During Takeoff- Climbout Operations.

LWP No. 211, April 13, 1966.

2. Copeland, W. L.; Hilton, D. A.; Huckel, V.; Dibble, A. C., Jr.; and Maglieri, D. J.: Noise Measurement Evaluations of Various Takeoff-Climbout Profiles of a Four-Engine Turbojet Transport Airplane.

NASA TN D-3715, 1966.

3.

Copeland, W. L.: Noise Measurements During Takeoff Climbout Operations of Four Jet Transports.

Presented at 74th Meeting of the Acoustical Society of America, Miami Beach, Florida, November 1967.

4. Copeland, W. L.; and Clark, L. R.

: Noise Measurements for a Three-Engine Turbofan Transport Airplane During Climbout and Landing Approach Operations.

NASA TN D-6137, 1971.

5. Schaefer, W.

T., Jr.; and Copeland, W.

L.: Noise Abatement Approach Flight Test Investigation of a Convair 880M Airplane.

LWP No. 305, 1966.

6. Zaloveik, J. A.: Effect of Thrust and Altitude in Steep Approaches on Ground Track Noise. NASA TN D-4241, 1967.

7. Sawyer, R. H.; and Schaefer, W. T.: Operational Limitations in Flying Noise Abatement Approaches.

NASA TN D-5497, 1969.

8. Quigley, H. C.; Snyder, C. T.; Fry, E. G.; Power, L. J.; and Innis, R. C.: Flight and Simulation Investigation of Methods for Implementing Noise Abatement Landing Approaches. NASA TN D-5781, 1970.

9. Meyersburg, R. B.; and Williams, C. H.: The Two-Segment Noise Abatement Approach to Landing. Presented at the International Conference on the Reduction of Noise and Disturbance Caused by Civil Aircraft, London, England, November 1966.

10.

Chubboy, R. A.: An Operational Evaluation of the Two-Segment Approach for Noise Abatement. FAA Report No. FAA-RD-71-72, April 17, 1972.

11.

Denery, D. G.; Bourquin, K. R.; White, K. C.; and Drinkwater, F. J., III: Flight Evaluation of Three-Dimensional Area Navigation for Jet Transport Noise Abatement. AIAA Paper No. 72-814, 1972.

IV - MILITARY

ASPECTS

In discussing the military aspects of noise technology, NASA's intent is specifically not to review or comment on the military's noise research and development programs.

Rather, it is to present the major cooperative programs underway between the DOD and NASA and to point out NASA's objectives and interest in participating jointly with the military services.

In general, at least three major benefits accrue to the Nation from combined efforts of this type: (1) NASA noise technology applicable to and needed by military programs is avail- able in a timely way.

(2) Civilian derivatives of military machines, environmentally acceptable with re- gard to noise and pollution, may be more economically and expeditiously phased into the commercial field.

(3) NASA keeps the touchstone of practical hardware in its long range technology ef- fort.

TURBINE ENGINE PROGRAMS The Air Force is in the process of procuring a demonstration model of an Advanced Technology Engine (ATE) to insure that the technology of high performance, high bypass ratio engines is sufficient and ready for application, this decade and the next, to sub- sonic military aircraft.

The engine generally is in the 20, 000-pound plus thrust cate- gory, will have a high thrust to weight ratio, good specific fuel consumption and, with minimum modification, contain those qualities attractive to civil propulsive systems.

The potential for the application of a derivative of this engine to commercial transport is significant, and the Air Force and NASA have recognized that a cooperative effort between the two agencies is desirable and necessary.

Accordingly, a close relationship, formally documented, exists between NASA and the Air Force during this program de- velopment to insure that specific features are designed into the engine to meet environ- mental requirements.

The Air Force obviously cannot compromise engine performance unduly to meet civil standards.

However, acceptable compromises and/or trade-offs in performance are being considered along with design features that minimize the modifica- tions that would be needed to provide an environmentally acceptable ATE derivative en- gine.

These considerations are important and recognized factors in the Air Force pro- gram.

the Advanced Technology Engine are in- demonstrator programs such as Technology on credible, full-scale ad- and develop specific engine systems based tended to explore military application for the ATE include both vanced technology hardware. Potential other support aircraft. The logistics transports and perhaps short-haul and long-haul high 20,000- to 25,000-pounds thrust, moderately cycle and design characteristics are improved maintainability through bypass ratio (~7), low specific fuel consumption, and emissions. The per- maintenance concepts, and reduced bare engine noise modular for civil con- cycle makes it particularly attractive formance expected from this engine objective in working with the takeoff and landing aircraft. NASA's primary ventional minimizing the environmental the establishment of those design features for Air Force is versions. Based on its both in the military and civilian impact of the operational engine predict the perfor- and technology programs, NASA can help propulsion noise research design with low noise and emissions and determine which mance tradeoffs associated ATE demonstrator interest, particularly with regard to the features are in the national civil derivative potential.

performance jet transport air- Powered lift technology for application to civil high Quiet, clean an important objective of NASA's propulsion research.

craft constitutes the more systems for this application present an even greater challenge than propulsion the very stringent noise and pollu- engine like ATE, primarily because of conventional areas.

with expected operation close to heavily populated tion requirements associated A very important part of NASA's powered lift technology program is the Quiet, Clean, with the ATE, this en- Short-Haul Experimental Engine (QCSEE) program. Compared a much higher bypass ratio (15), may have variable rather than gine will probably have and will have very fixed pitch fan blades, may be geared rather than have direct drive, velocities (700 ft/sec) to meet the low noise goals for civilpoweredlift aircraft.

low exit turbine spool differs greatly from the ATE, it is quite Although the fan/low pressure that the engine core can be common to the two programs, perhaps with rela- possible As the two programs progress, every effort will be made to tively minor modification.

ATE core and to apply it to the QCSEE pro- capitalize on the military development of the will begin with that intent and gram to reduce risks and costs. The QCSEE program will continue unless that course proves technically impossible.

STOL AIRCRAFT PROGRAMS STOL (AMST) program is being conducted by the Air Force The Advanced Medium the worth of to design and fabricate two different STOL airplanes to test and evaluate aircraft in a modernized tactical airlift fleet. NASA is actively supporting such this program to generate quiet propulsive lift technology and design data for civil STOL, combined program are RTOL, and CTOL transport. Prime NASA objectives in this directed toward high angle takeoff and approach paths (permitted through the use of pro- pulsive lift) to provide community noise reduction and relief of airport congestion (both air and ground movement of aircraft).

Further, NASA's research using this military aircraft will establish a technical foundation for the certification, operation, and regula- tion of civil propulsive lift transports.

Complete satisfaction of these significantly different objectives imposes diverse vehicle requirements that clearly would result in completely different aircraft designs.

However, a very significant portion of the NASA objectives can be satisfied through this program.

For example, propulsive lift flight research requires variations in noise re- duction, lift-coefficient, thrust-weight ratio, aircraft loadings, stability and control systems, and high lift systems considerably broader than those achievable through the point-design prototypes.

Nevertheless, these prototypes can provide the baseline tech- nical and operational data within certain ranges of parametric interest. NASA participa- tion in this program is aimed not only at supporting the military objectives wherever possible but also at using this aircraft in a test program to demonstrate the viability of a civil STOL transport that is environmentally acceptable to surrounding communities.

ROTORCRAFT PROGRAMS Two research programs, being conducted jointly with the Army, have a direct bear- ing on noise abatement and noise research.

The first, the Tilt Rotor Research Aircraft embodies the desirable hover charac- teristics of the turboprop aircraft through tiltable rotors mounted at the wing tips.

The low disk loading r'otors, the absence of a tail rotor, the reduced power required because of wing lift, and the more efficient rotor performance in the cruise configuration com- bine to offer a substantial noise reduction potential. A goal of 90 perceived noise deci- bels (PNdB) along a sideline distance of 500 feet (ft) for the most critical flight condition has been established.

The Army is concerned with cruise flight noise and the resulting detection times.

Comparison of the noise of the tilt-rotor aircraft with that of a single rotor helicopter, both in cruise and at low altitude, promises a reduction from around 125 to about 65 PNdB with the corresponding detection time reduced from that of 36 to 72 seconds to about 5 to 10 seconds.

Objectives of the program include: (1) The exploration, through flight research, of current technology of interest to the aircraft community for the development of useful, quiet, easily maintainable commercial or military V/STOL tilt-rotor aircraft, particularly by verifying rotor/plyon/wing dynamic stability and aircraft performance over the entire envelope speed on downwash and rotor disk and tip of the effects of tilt (2) The investigation on hover mode operations.

noise and the impact and evaluation of methods the development and objectives include Other research and noise and to to reduce congestion operation for efficient near-terminal procedures aircraft of applying tilt rotor of potential benefits and the evaluation increase safety, Army Air Mobility missions.

to various capabilities and is being designed the second program, Research Aircraft, The Rotor Systems essentially be a com- bed. The aircraft will as a flight research test developed to serve testing throughout propulsion to permit rotor with a wing and auxiliary pound helicopter of operating conditions.

a wide range and high-speed for- low-speed flight, missions entail hovering, Today's helicopter flow condi- flight creates asymmetrical of a rotor in forward ward flight. The operation and/or advancing blade compressibility resulting in retreating blade stall tions finally loads are a source of vibratory unsteady aerodynamic flows effects. These generally in various stages of are, however, new rotor concepts aerodynamic noise. There and of these limitations, thereby the potential of alleviating some development that have as a military vehi- transportation system and the helicopter as a commercial enhancing the blades azimuthal distribution of geometry rotor offers a favorable cle. The variable blade planform designs are vibration. New tip shapes and that may reduce noise and efficiency and reduced noise.

potential improvements in aerodynamic considered as for the investigation of the rotor Research Aircraft will be used The Rotor Systems The design of the vehicle demonstrate low-noise characteristics.

system intended to inlet nacelles and exhaust ducts, will consider noise suppression equipment on engine masking of the rotor system device in order to prevent gear boxes, and the antitorque An external noise level other aircraft systems and subsystems.

noise with noise from in takeoff and landing and, excluding more than 95 PNdB at the 500 ft sideline of not 500 ft from the aircraft when not more than 89 PNdB at any point main rotor noise, to 150 knots has been established.

helicopter at speeds up operating as a conventional 5-11 Office: 1973- 758-034/5228 Region *U.S.Government Printing