NASA-CR-114761-1973 Component noise variables of a light observation helicopter《轻型观察直升机的部件噪声变量》.pdf

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1、-NASA CONTRACTORREPORT NASA 6C /a76/HH 73-41 -9/L L /LCOMPONENT NOISE VARIABLES OF ALIGHT OBSERVATION HELICOPTERBy: Frank Robinson(NASA-CE-11 Ib) COMPODET iOISE N74-20652VA:IABLES OF A LIGHT OBSERVATIO1BELICCPfER (Huqhes Helicopters, CulverCity, Calif.) 78 p HC $7.50 CSCL 01C UnclasG3/02 35586Prepar

2、ed Under Contract No. NAS 2-7254By: HUGHES HELICOPTERSDivision of Summa CorporationCulver City, CaliforniaFor: Headquarters, U. S. Army Air Mobility R;,; : :; ;I F/-40Figue 2.“Quet“ eliopte in BaslineLes Engne“ onfguraionEngneExaut s ilnce b Lrg TnkMufle a EtrmeRihtProvided by IHSNot for ResaleNo re

3、production or networking permitted without license from IHS-,-,-When mounted on the test rig, the helicopter could be run with any combinationof its major components either removed or silenced. The tail rotor could beremoved and the engine silenced so only the main rotor could be heard. Themain roto

4、r could be removed and the engine silenced so only the tail rotorcould be heard. Both the main and tail rotors could be removed with thedynamometer absorbing the power so only the engine could be heard. Anoise level difference of 5 or 10 decibels in the frequency range of interestbetween the silence

5、d components and the components being investigated isusually sufficient to make their noise contribution negligible.The test rig worked exceptionally well. The dynamometer was able to absorbfull engine power. Its cooling system dissipated the rejected heat withoutany problems and the noise level of

6、the cooling fans was only 68 decibels.Narrow band spectra plots for the cooling system are presented as runs 35and 238 in Appendix I. A narrow band spectra plot of a typical ambient isshown after run 238. There was no evidence of ground resonance or anyother dynamic instability. There was, however,

7、a noise frequency recordedwhich corresponded to the RPM of the dynamometer drive shaft. This couldpossibly be corrected by improving the shaft balance. This test rig shouldprovide a convenient tool during future tests, for investigating the noise fromany isolated component.Test ProcedureThe test air

8、craft was equipped with precision visual instrumentation forreading engine torque, tail rotor torque, tail rotor thrust, collective pitchand tail rotor pitch. The aircraft was flown in free hover at a 6-foot skidheight and a variety of gross weights and rotor speeds to obtain calibratedreadings. Thi

9、s enabled the pilot to duplicate the various rotor thrust andpower conditions with the helicopter mounted on the test rig by setting-upthe same values for collective pitch, etc., as those recorded during free hover.The noise tests were conducted between midnight and approximately 5:00 AMto obtain th

10、e lowest background noise and the calmest wind conditions. Theacceptable winds were limited to three knots and no visible precipitation waspermitted. The relative humidity and temperature were monitored and alltests were conducted under similar ambient conditions. Also, tests weredelayed whenever ai

11、rcraft were observed flying anywhere in the surroundingarea. The test rig was located in an open area where there was a minimumof reflective surfaces. The control microphone (Position No. 1) was located200 feet from the helicopter at an azimuth position 30 degrees left of due aft(Figure ZA). The mic

12、rophone at Position No. 2 was located at 200 feet,30 degrees left of forward. The microphones were four feet above the groundand the terrain between the helicopter and the microphones was primarilygrass which reduced the effects of ground reflection waves.Provided by IHSNot for ResaleNo reproduction

13、 or networking permitted without license from IHS-,-,-200 FTMICROPHONETEST RIG6 FT 4 FTFigure 2A. Microphone Height and Location Relative to Test HelicopterData Acquisition and ProcessingThe noise data was recorded at 60 inches-per-second on one-inch magnetictape. The data runs had a duration of 40

14、seconds each, with voice identifica-tion. All sound-pressure-level (SPL) data was referenced to 0. 0002 dynesper square centimeter.6Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The following equipment was used for recording the data:Position No. 1

15、 - B&K Type 4131 Microphone, B & K Type 2203Sound Level MeterPosition No. 2 - B &K Type 4145 Microphone, B & K Type 2107Soufid Level MeterAcoustic Calibrator - B & K Piston Phone Type 4220Tape Recorder - Pemco Model 120The data recorded at Position No. 1 was then processed by Wyle Laboratoriesusing

16、a real-time-analyzer and a digital computer program for calculatingperceived-noise-level (PNdB). In addition to the calculated PNdB value andthe linear readings, values were also recorded using “A“ scale and “D“ scaleelectrical weighting networks. Each data point was based on an average valueobtaine

17、d using a 20 second portion of the record. In addition to the tabulateddata, selected one-third octave plots and narrow band plots were made for avariety of test configurations.Limitations of Test DataRange of instrumentation. - The equipment used for recording sound pressurelevels has a limited ran

18、ge of about 40 or 45 decibels. For each sound levelbeing recorded, the operator set his equipment to record at the linear overall-sound-pressure-level (OASPL) indicated by his visual reading. When makinga frequency spectra plot or when computingthe perceived-noise-level (PNdB),if some of the sound p

19、ressure levels were considerably below the maximumlevel, they would become mixed with the instrumentation noise. During thistest program, the high frequency (5, 000 to 20, 000 Hz) helicopter noise levelswere usually quite low. On many, if not most, of the one-third octave plotscontained in this repo

20、rt, the high-frequency portion of the spectra is actuallyinstrumentation noise and moves up or down in amplitude as the operatoradjusts the range setting of the recorder. The recording level is shown inparenthesis on each plot to aid in evaluating the data.In future tests, when it is desired to impr

21、ove the accuracy and resolution ofthe low-level, high-frequency noise sources, it is recommended that separateadditional records be made with low range settings on the recording instru-mentation. These records can then be used in conjunction with the initialhigh-level recordings to cover the entire

22、sound pressure level and frequencyrange of the helicopter.7Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Ground reflection waves. - Frequently, when measuring sound pressure, themicrophone will record two (or more) sound pressure waves coming from

23、thesame source. One pressure wave travels direct (line-of-sight), while theother wave is first reflected off the ground, or other surfa-ce, and thentravels to the microphone. Since the reflected wave must travel a greaterdistance, it will arrive at the microphone some time increment after thedirect

24、wave. This will produce a phase shift between the two waves and thusthe reflected wave may either augment or diminish the direct wave, dependingon the resulting phase shift.The possible influence of ground reflection waves on the data contained inthis report is discussed in Appendix II. In general,

25、their effects did notseriously impair the overall quality of the data.TEST RESULTSComplete Helicopter vs Component NoiseA wide variety of helicopter configurations were tested. These are listed intables contained in the appendix of this report. The recorded noise levelsof the three major components

26、of the OH-6A helicopter are listed below.Records taken at four gross-weight test conditions were averaged to obtainthese values and all are at 103 percent N2 engine speed. The three separatecomponent noise levels were added together, using the method described inreference 3, and are compared with th

27、e levels recorded for the completehelicopter. The agreement between the sum of the component noise and thenoise of the complete helicopter, and, the agreement between the completehelicopter in free hover and when mounted on the test rig, are quite good.Linear “D“db Weighted PNdBMain Rotor Only 85 74

28、 79Tail Rotor Only 86 82 89Engine Only 82 79 86Sum of Components 90 84 91Complete OH-6A, Mountedon Test Rig 89 85 92Complete OH-6A, In FreeHover 87 84 918Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Main Rotor NoiseAs discussed in Appendix II, the

29、 height of the main rotor tended to produceground reflection waves which distorted the sound pressure readings, partic-ularly from 500 to 1000 Hz. It is at this frequency that the broad band noisefrom the main rotor has its greatest influence on the calculated perceived-noise-level (PNdB). Since the

30、 linear overall-sound-pres sure-levels (OASPL)were less affected by ground reflection waves, they were used to develop theparametric curves for the main rotor noise.Tip speed. - Figure 3 shows the variation of OASPL with tip speed for themain rotor of the.standard OH-6A helicopter. These tests were

31、conductedwith the tail rotor removed and the engine silenced so that the dominantnoise source was the main rotor. Figure 4 shows the OASPL vs tip speedvariation for the “Quiet“ helicopter 5-bladed main rotor with tapered tips.880O86848280( 90060 RUN 117_J0 1050w 7070-Jw 60 RUN 69w60 RUN 11220 50 100

32、 200 500 1000 2000FREQUENCY- HZFigure 7. 1/3 Octave Spectra Comparison of4-Bladed Tail Rotors with DifferentBlade Azimuth Spacing14Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Blade azimuth spacing. - With a 4-bladed tail rotor, the azimuth spacin

33、gbetween blades can be at angles other than ninety degrees by displacing onepair of blades in a scissor-like motion relative to the other pair. By arrang-ing the blades at other than ninety degrees, the 4/rev noise component canbe reduced. However, as the 4/rev is reduced the 2/rev is increased. Thi

34、sis apparent by comparing the 600-1200 tail rotor with the 900-900 tail rotorin Figure 7.There is also an increase in the 3/rev and 6/rev with the 600-1200 rotor.A 6/rev would be anticipated but it is difficult to visualize the origin of a3 /rev. Coincidentally, the engine drive shaft to the dynamom

35、eter alsorotates at 103 RPS and its universal joints would produce a 206 Hz frequency.This could be the noise source at these frequencies, rather than the tail rotor.Again using the average noise level for five combinations of thrust and tipspeed, the noise level is compared for three blade spacings

36、 The 750- 1050spacing appear to be the quietest and the 600-1200 the noisiest, with the900-900 falling in between. However, the differences are small and the lowtip speed of the “Quiet“ helicopter tail rotor (495 feet per second) places itsnoise level only slightly above that of the silenced engine,

37、 making the accuracyof these measurements about the same as the differences measured betweentail rotors.Linear Different Differentdb from 900 PNdB from 900600 - 1200 Blade Spacing 74.8 +1.0 78.9 +2.4750 - 1050 Blade Spacing 72.2 -1.6 75.0 -1.5900 - 900 Blade Spacing 73.8 -0- 76.5 -0-15Provided by IH

38、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Blade construction and/or airfoil. - Tail rotor blades fabricated of Fiber-glas and steel were compared with blades fabricated of aluminum. Also,blades with a symmetrical airfoil (NACA 0014) were compared with blades

39、having a cambered airfoil (NACA 63-415 MOD). The cambered aluminumblades had a chord length of 5. 3 inches while the symmetrical Fiberglas andthe cambered Fiberglas blades had a chord length of 4.8 inches. Otherwise,the rotors and test conditions were identical.All three rotors had noise levels with

40、in a total spread of one decibel, usingPNdB, linear, or “D“ weighted values. This would be remarkable agree-ment even for the same rotor during two different test runs. Thus it can beconcluded that all three rotors had essentially the same noise level. Thesetests were conducted using the standard OH

41、-6A helicopter with everythingsilenced except the tail rotor. The higher tip speed of the OH-6A tail rotor(692 FPS) put the tail rotor noise level well above any other noise source,making this data quite reliable.Linear “D“db Weighted PNdBCambered Aluminum Blades 83.8 78.9 86.0Cambered Fiberglas & S

42、teel Blades 82. 9 78. 1 85.6Symmetrical Fiberglas & Steel Blades 83.8 78.7 86.0Thrust. - Figure 8 is a plot of tail rotor thrust versus noise at a constanttip speed and with constant blade area. This again means that the averagelift coefficient of the blade sections must increase as the thrust incre

43、ases.In a new design, the blade area would be increased to retain the same maxi-mum thrust capability which could have some additional effect on the noiselevel variation with thrust.Powerplant NoiseBoth aircraft had basically the same powerplant. However, the engine in the“Quiet “ helicopter did hav

44、e a number of minor factory modifications to lowerits noise level. These modifications are described in detail in Reference 1.Figure 9 shows the noise level variation with power of the standard OH-6Apowerplant. Narrow band spectra plots are included in Appendix I for the“Quiet“ powerplant (Run 30) a

45、nd for the standard OH-6A powerplant (Run 201).A narrow-band plot is also included for the OH-6A powerplant with the inletsilenced, insulated cowl doors installed, and exhaust silenced (Run 180).16Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-9492

46、88-J86a NOISE db84 = .041aTHRUST LB 0 40 80 120 160 200 240THRUST- LBFigure 8. Variation in Tail Rotor NoiseLevel With Thrust90a NOISE db.046 .488 - - a POWER HP86 PNdb84 H TLINEAR8280 NI08. SCALE7876160 170 180 190 200 210 220 230 240 250 260 270ENGINE POWER - HPFigure 9. Noise Variation With Power

47、 ofStandard OH-6A Powerplant17Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Two prominent frequencies occur on these plots, at 103 Hz and 206 Hz.These are the first and second harmonics of the engine output drive shaftfrequency. Any out-of-balance

48、of the drive shaft or misalignment of itsuniversal joints, would tend to vibrate the aircraft at these frequencies. Itis apparent from the narrow-band plot of Run 243, “Quiet“ main rotor only,that the 103 Hz frequency does not exist when the dynamometer is removed.It is difficult to tell on the othe

49、r spectra plots because both the tail rotor andthe 4-bladed main rotor have harmonics at those frequencies. The existenceof these two noise sources should not invalidate the comparative data, how-ever, since the same shaft was used on all runs.Exhaust muffler and insulated cowl doors. - Figure 10 shows third-octavespectra plot