Document
NAS-. im3nical Memorandum 85850 USAAV RADCOM Technical Report-83-A-14
1cl"lodelHelicopter Rotor High -Speed
Impulsive Noise: Measured
Acoustics and Blade Pressures
D.A. Boxweil, F.H. Schmitz, W.R. Splettstoesser
and K.J. Schultz
September 1983 United States Army Aviation Fiesearch and Development National Aeronautics and Space Administra:ion Command NASA Technical Memorandum 85850 USAAV RADCOM Technical Report-83-A-14
Model Helicopter Rotor High =Speed
Impulsive Noise: Measured
Acoustics and Blade Pressures
D. A. Boxwell, F. H. Schmitz, Aeromechanics Laboratory, Research and Technology Laboratories, U. S. Army AviLtion Research and Dpvelopment Command, Ames Research Center, Moffett Field, California W. R. Splettstoesser, DFV LR, lnstitut f. Entwurfsaerodynamik, K. J. Schultz, Abt. Techrische Akustik, Brjunschweig, W. Germany National AerGnautics and United States Army Space Administration Aviation Research and Development Command Ames Research Center Moffett Field California 94C35 St LOUIS, Missouri 63166 M O D E L HELICOPTER ROTOR HIGH-SPEED WULSIVE NOISE: MEASURED ACOUSTICS -WD BLADE PRESSURES D. A. Boxweil and F. H. Schmitz Aeromechanics Laboratory U . S . Army Research and Technology Labcratories N A S A Ames Research Center Moffett F i e l d , C a l i f o r n i s 94035 U.S.A.
W. R. S p l e t t s t o e s s e r and K. J. Schultz
DFVLR, I n s t i t u t f . EntlJurf saerodynarcik
Abt. Technische Akustik, Braunschweig, W . Germany Abstract A 117-scale research model of the AH-1 series h e l i c o p t e r main r o t o r was t e s t e d i n the open-jet anechoic test s e c t i o n of t h e Deutsch-
Niederlaendischer Windkanal (DNW) (Netherlands) . Model-rotor a c o u s t i c
and sinultaneous blade pressure data were recorded a t high forward speeds where f u l l - s c a l e h e l i c o p t e r high-speed impulsive noise levels are h o w n t o be dominant. Model-rotor measurements of the peak acous- t i c pressure l e v e l s , waveform shapes, and d i r e c t i v i t y p a t t e r n s are d i r e c t l y compared with f u l l - s c a l e i n v e s t i g a t i o n s , using an equivalent in-flight technique. Model a c o u s t i c data are shown t o scale remark- ably w e l l i n shape and i n amplitude with f u l l - s c a l e r e s u l t s .
Model rotor-blade pressures are presented f o r r o t o r operating conditions both with and without shock-like d i s c o n t i n u i t i e s i n the r a d i a t e d acoustic yaveform. Acoustically, both model and f u l l - s c a l e measure- ments support c u r r e n t evidence t h a t above c e r t a i n high subsonic advancing-tip Mach numbers (MA 1 0.9), l o c a l shock waves t h a t e x i s t T on the r o t o r blades "delocalize" and r n d i a t e t o t h e a c o u s t i c f a r - f i e l d .
1. Introduction Over the p a s t few years, the understanding and p r e d i c t i o n of high-speed imp1Jlsive noise of h e l i c o p t e r r o t o r s have improved dramati- Through fundamental experiments i n an aerodynamically and c a l l y .
a c o u s t i c a l l y cviitrolled hover environment and thrcugh p r e c i s e numeri- c a l computation, some of the important parameters of t h i s annoying and detectable r o t o r noise have been i s o i a t e d . These fundamental s t u d i e s have shown t h a t i n a d d i t i o n t o blade geometry, the l o c a l transonic flow f i e l d s u r r o w d i n g the transonic hovering r o t o r can be an important and v i t a l contributor t o the radiated noise (Refs. [l-31).
A t l o c a l l y transonic conditions, shocks near the t i p of the f i n i t e - thick r o t o r blade can escape ("delrxalize") t o the acoustic f a r - f i e l d The t i p Yach mmber a t which t h i s phenomenon is observed (Ref. [ l ] ) .
i s c a l l e d the "delocalization Mach number" and is s t r o n g l y a f f e c t e d by the l o c a l blade geometry i n the t i p region.
17-1 Although the hovering r o t o r is i d e a l f o r the fundamental study of transonic aerodynamics and noise, it does not exactly duplicate t h e flow environment of h e l i c o p t e r r o t o r s which normally r a d i a t e high- speed Impulsive noise. The primary reason i s the l a r g e v e l o c i t y asymmetry which e x i s t s on h e l i c o p t e r s i n high-speed forward f l i g h t .
On t h e advancing s i d e of the r o t o r , t h e rotor-blade r o t a t i o n a l veloc- i t y sums with the h e l i c o p t e r ' s forward v e l o c i t v , o f t e n causing advancing-tip speeds t o approach Mach 1. On the r e t r e a t i n g s i d e of the r o t o r , the r o t a t i o n a l v e l o c i t y and forward v e l o c i t y are i n oppo- site d i r e c t i o n s y i e l d i n g low e f f e c t i v e v e l o c i t i e s . A s a r e s u l t , the h e l i c o p t e r r o t o r blade sees a c c n t i n u a l l y changing v e l o c i t y f i e l d as i t r o t a t e s through each revolution. On the advancing s i d e of the r o t o r , where v e l o c i t i e s are high, angles of a t t a c k are small; on the r e t r e a t i n g s i d e , the angles 3f a t t a c k are q u i t e high because o f low relatJ.ve v e l o c i t i e s . Therefore, i n high-speed forward f l i g h t , rotor- blade dynamic stall on the r e t r e a t i n g s i d e of t h e r o t o r is a real p o s s f t i l i t y . To avoid t h i s problem, the h e l i c o p t e r rotor-blade must not be too t h i n (0.06 t o 0.12 c ) , a c o n s t r a i n t t h a t aggravates t h e advancing-blade transonic aerodynamic and high-speed noise problems.
are high, the r e l a t i v e l y t h i c k a i r f o i l s I f advancing-tip Mach numbers encounter l a r g e regions of unsteady transonic flow s d d a l o c a l i t a t f a n occurs, allowing l o c a l shock waves on the a i r f o i l surface t o r a d i a t e to the acoustic f a r - f i e l d (Ref. [4]).
Acoustic p r e d i c t i o n of the r e s u l t i n g noise f o r high-speed h e l i - copter f l i g h t was attempted some pears ago using l i n e a r models with limited success ( R e f . [ 5 ] ) . A s i n the high-speed hover problem, non- l i n e a r transonic aerodynamics play a s i g n i f i c a n t r o l e i n the a c o u s t i c r a d i a t i o n and must be included i n t h e modeling. However, the high- speed h e l i c o p t e r transonic flow f i e l d is unsteady. This f u r t h e r complicates the aerodynamic transonic f low-f i e l d c o q m t a t i o n s by making f ar-f i e l d acoustic p r e d i c t i o n s of high-speed impulsive noise more d i f f i c u l t than the transonic-hover problem. To date, no computer code has successfully predicted the amplitude o r waveform of high- speed impulsive noise i n forward f l i g h t when shock waves d e l o c a l i z e t o the acoustic f a r - f i e l d .
An a l t e r n a t i v e t o the a n a l y t i c a l approach is to d u p l i c a t e the f u l l - s c a l e high-speed impulsive noise i n model scale. Although t h i s procedure seems straightforward, most previoiis a c o u s t i c f u l l - s c a l e , ground-based, and wind-tunnel t e s t i n g required data c o r r e c t i o n s t h a t made q u a n t i t a t i v e comparisons q u i t e d i f f i c u l t . Early model-to-full- scale comparisons on a hovering r o t o r were encouraging b u t q u a l i t a t i v e
(Ref. 16 j ) . More recent model-scale wind-tunnel comparisons with
ground-based, f u l l - s c a l e measurements (Ref. [ 7 ] ) showed general agree- ment a f t e r the data were corrected f o r mic.ophone p o s i t i o n d i f f e r e n c e s , reverberation, atmospheric and ground a t t e n u a t i o n , and forward veloc- i t y e f f e c t s (Doppler s h i f t ) .
Many of these data c o r r e c t i o n s a r e not necessary i f the f u l l - s c a l e data a r e gathered by using the i n - f l i g h t technique developed by the Aeromechanics Laboratory (Qef. [SI). The technique c o n s i s t s of mounting microphones on a q u i e t , fixed-wing a i r c r a f t t h a t is flown in f o r m t i o n with the subject h e l i c o p t e r over a matrix of f l i g h t condi- t i o n s . Because relaLive geometry between the measurement a i r c r a f t 17-2 and subject h e l i c o p t e r is e s s e n t i a l l y f fxed, t h e procedure is equiva- l e n t t o wind-tunnel testing. I n f u l l - s c a l e f l i g h t , t h e microphone and subject h e l i c o p t e r maintain t h e same r e l a t i v e l o c a t i o n s while f l y i n g through a i r a t t h e same f l i g h t conditions. I n wind-tunnel t e s t i n g , the in-flow microphones and t h e --otor i n a uniform a i r flow a r e fixed i n space a t s i m i l a r relative positions. Aerodynamically and a c o u s t i c a l l y the procedures are t h e same, provided t h a t t h e back- ground noise and acoustic r e f l e c t i o n s of both i n - f l i g h t and wind- tunnel t e s t i n g are minimized.
The in-flight procedure w a s used t o gather a c o u s t i c data over a f a i r l y complete matrix of f l i g h t conditions f o r t h e UH-1H helicop- ter (Ref. [ a ] ) . Because t h i s w a s r e a l l y t h e f i r s t series of tests with the i n - f l i g h t , f a r - f i e l d measuring technique, some a c o u s t i c data scatter w a s evident. These data were f i r s t compared with 1/7-scale model UH-1H acoustic data gathered i n an a c o u s t i c a l l y t r e a t e d , hard- walled wind tunnel (Ref. [ 4 ] ) . Although tunnel reverberation e f f e c t s did influecc the r e s u l t s of the model-scale data a t lower adrancing- t i p Mach iiumbers, reasonably good q u a n t i t a t i v e model-scale t o f u l l - scale c o r r e l a t i o n w a s achieved. Certainly, many of t h e f u l l - s c a l e teaporal shapes and amplitude trends were duplicated i n t h e model- scale experiment .
Another attempt a t model-to-full-scale r o t o r a c o u s t i c s c a l i n g was made with the AH-1 s e r i e s helicopter. Carefully controlled per- formance and i n - f l i g h t acoustic tests ( R e f s . [9] and [ l o ] ) were run on a standard AH-1/540 r o t o r over a f u l l matrix of f l i g h t conditions
and compared with a l/i-scale model operational-loads-survey (Om)
rotor tested under similar conditions i n a 3-m (9.8-ft) open-jet,
anechoic wind tunnel , CEPR-19 (Ref. [ 111 ) . Although t h e blade-vortex
i n t e r a c t i o n impulsive noise phenomenon w a s shown t o scale f a i r l y w e l l , wind-tunnel speed c o n s t r a i n t s did not allow t h e high-speed impulsive noise scaling question t o Le f u l l y addressed.
This same OLS model r o t o r was r e c e n t l y t e s t e d i n t h e acousti- c a l l y t r e a t e d DNW over a f u l l range of t e s t i n g conditions, some of which are reported here. The emphasis i n t h i s paper is on c a r e f u l l y exploring the important scaling parameters of helicopter-rotor, high- s p e e d impulsive noise. I n addition, the d i r e c t i v i t y of t h e r a d i a t i n g noise f i e l d i s documented along with selected blade pressures a t sev- e r a l high subsonic advancing-tip Mach numbers.
2. Experimental Des* The model-rotor test was performed i n the new Cerman-Dutch wind tunnel, Deutsch-Niederiaendischer Windkanal (DNW), which is located about 150 k m northeast of Amsterdam, the Netherlands; an aerial view of t h i s f a c i l i t y i s shown i n Fig. 1. (The t e s t i n g and subsequent d a t a reduction e i f o r t s a r e part of a continuing memorandum of understanding €or cooperative research between the German and United S t a t e s govern- ments.) The DNW i s a subsordc, atmospheric wind tunnel of t h e closed- return t y p e ; i t has three interchangeable, closed, t.est--section con- f i g u r a t i o n s and one open-jet coilfiguration with a 6- by 8-m (19.7- by 26.2-ft) contrac.tion. The tunnel was designed €or low bac!<ground 17-3 was used on t h e t h r e e main-microphone support s t r u t s located w i t h i n t h e f r e e j e t . I n the i n i t i a l s t a g e s of t e s t i n g , foam treatment of the microphone supports w a s n o t thought t o be necessary. However, on-line a c o u s t i c c a l i b r a t i o n of t h e test configuration showed r e f l e c - t i o n s from the support s t r u t s , which would have d i s t o r t e d the measured r o t o r a c o u s t i c s i g n a l s (see Appendix B of Ref. [12]). Additional a c o u s t i c c a l i b r a t i o n s with and without flow yielded t h e f i p a l test configuration shown i n Fig. 2.
A t o t a l of 19 B&K 1/4-in. microphones (Type 4135) were dis- t r i b u t e d around the r o t o r , 10 i n t h e open-jet core flow and 9 o u t of The l o c a t i o n of t h e in-flow microphones, t y p i c a l l y 3.26 m it (Fig. 3 ) .
(10.7 f t ) from the r o t o r hub, w a s chosen t o correspond t o a n average scaled microphone p o s i t i o n of t h e f u l l - s c a l e a c o u s t i c tests of Refs. [9] and [ l o ] . A l l of the in-flow microphones (Nos. 1-4, 6-10, and 15) were positioned forward and down from t h e rotor-hub plane, where severe impulsive noise is known t o r a 2 i a t e (Ref. [SI). Another in-flow microphone (No. 1) w a s placed in-plane d i r e c t l y i n f r o n t of t h e nozzle l i p a t exactly twice the d i s t a n c e (6.52 m (21.4 f t . ) ) .
In-plane forward f l i g h t a c o u s t i c decay rates were measured by compar- ing a c o u s t i c l e v e l s from microphones Nos. 1 (6.52 m) and 2 (3.26 m).
One a d d i t i o n a l microphor,a (No. 15) w a s placed i n t h e flow underneath the r o t o r i n the model-scaled cabin p o s i t i o n t o measure impulsive noise t h a t would be heard i n t h e cabin. The in-flow microphones were c a r e f u l l y foam-bedded i n s p e c i a l a d a p t e r s and equipped with nose cones pointing upstream i n t o the oncoming flow. The out-of-flow micro- phones - most of which were located near t h e f l o o r (Nos. 5, 11, 16-19), above the r o t o r plane (Nos. 1 2 and 13). and toward t h e a f t quadrant (No. 14) of the r o t o r - were used t o gather a d d i t i o n a l information about d i r e c t i v i t y , d i s t a n c e , and shear-layer e f f e c t s . They were equipped with standard g r i d and wind screens and o r i e n t e d f o r grazing incidence. Microphone c a l i b r a t i o n s were accomplished with B&K piston- For phones on the beginning o r end of each recorded magnetic tape.
intermediate checks, the i n s e r t voltage method was applied. Although not an absolute c a l i b r a t i o n method (because the diaphragm s e n s i t i v i t y is not included i n the c a l i b r a t i o n procedure), i n s e r t voltage can be e a s i l y applied a s an e l e c t r i c a l checkout of t h e microphone c i r c u i t .
The importance of having "simple" a c o u s t i c "checks" i n a l a r g e t e s t i n g f a c i l i t y , such as t h e DNW, should be emphasized. These procedures helped t o a s s u r e t h e v a l i d i t y of the recorded d a t a and thereby mini- mize unproductive t e s t time. I n a d d i t i o n , c a l i b r a t i o n d a t a f o r t h e blade-pressure transducers were recorded a t 0 and 2.76 N/cm2 (0 and 4 l b / i n . 2 ) s t a t i c pressure. This W Z I ~ dune simultaneously f o r a l l transducers by placing the blade within a p o r t a b l e p l a s t i c c y l i n d r i - c a l sleeve and evacuating i t t o the desired c a l i b r a t i o n pressure.
The model-rotor blades were mounted on a teetering-hub assem- b l y , with the c o l l e c t i v e and the l o n g i t u d i n a l and lateral c y c l i c r o t o r i n p u t s provided by remotely c o n t r o l l e d e l e c t r i c swashplate a c t u a t o r s .
Tip-path-plane tilt was c o n t r o l l e d d i r e c t l y through c y c l i c , t h e s h a f t being r i g i d l y mounted i n t h e v e r t i c a l p o s i t i o n on the r o t o r stand. A six-component strain-gauge balance, comprising t h e top p o r t i o n of t h e drag, and r o t o r stand, was used t o monitor and record r o t o r t h r u s t , p i t c h i n g moments f o r each t e s t condition. Blade-surface pressures, blade flapping, and tunnel temperature were a l l measured i n t h e 17-5 0 !N PLANE 0 30" BELOW PLANE
n 450 BELOW PLANE
A NEARFLOOR t y Iml 0 CEILING 1 9 v CABIN POSITION (No. 15) / I / / / X Iml -U (a) Top view.
Fig. 3 . Microphone loca. ions in the DNW.
17-6 \ \ \ \ \ \ \ (b) Side view.
Fig. 3 . Concluded.
17-7 rotating-hub frame; measurement d a t a were transmitted by wires t o the lower end of t h e s h a f t , and t r a n s f e r r e d t o a nonrotating frame v i a a 156-channel s l i p - r i n g assembly. Rotor-shaft r o t a t i o n a l encoders of l/rev, 60/rev, and 180/rev were used t o accurately c o n t r o l rpm and t o azimuthally index r o t o r events. An automatic servo system w a s used t o c o n t r o l a variable-frequency, 90-hp electric motor t h a t drove t h e model r o t o r up t o 3,000 rpm. For a l l test cases, the ambient d r i v e system noise w a s below t h e wind-noise f l o o r , with p r a c t i c a l l y no e f f e c t on the r o t o r noise measurements (see Appendix B of Ref. [12]).
A l l the microphone s i g n a l s and s e l e c t e d pressure d a t a were monitored on-line and, a f t e r proper s i g n a l conditioning, simultaneously recorded on t h r e e multichannel, FM, magnetic tape recorders with a t o t a l of 60 channels set f o r a recording speed of 76.2 cm/sec (30 i p s ) and a frequency response of 20 kHz. Flapping potentiometer output, balance data, and a l s o t h e outputs from a hot-wire probe placed near micro- phone No. 6, and from one accelerometer positioned i n t h e blade near the t i p were recorded along with the a c o u s t i c and pressure data.
IRIG-B time-code and r o t o r azimuth s i g n a l s w e r e recorded on each tape f o r synchronization purposes. During the 1-min d a t a recording, two -3P 54208 FFT analyzers were used t o generate on-line instantaneous and averaged time-histories of s e l e c t e d microphones and pressure transducers. Wind-tunnel v e l o c i t y , temperature, and dew point, as w e l l as r o t o r speed, swashplate c o n t r o l inputs, and balance informa- t i o n , were processed on-line, using a portable HP 85 computer d i r e c t l y connected t o the r o t o r balance.
3. Model-Rotor C h a r a c t e r i s t i c s One of the main purposes of the test program w a s t o quanti- t a t i v e l y e s t a b l i s h the v a l i d i t y of using scaled model-rotors t o dupli- c a t e f u l l - s c a l e a c o u s t i c t e s t r e s u l t s . Therefore, the i n - f l i g h t data-gathering technique was used t o o b t a i n high-quality, s t a t i o n a r y , full-scP.le data on the AH-1 series h e l i c o p t z r (Cobra) with the BHT 540 r o t o r blades. These d a t a , gathered and :educed b!- t h e Aeromechanics Laboratory with the a s s i s t a n c e of the Army Aviation Engineering F l i g h t A c t i v i t y , Edwards A i r Force Base, and NASA Ames Research Center, i s presented i n d e t a i l i n Refs. [9] and [ l o ] . More r e c e n t l y , Aries Research Center has ran a second series of i n - f l i g h t a c o u s t i c tests on an AH-1G h e l i c o p t e r equipped with OLS r o t o r blades (Ref. [14]).
The thickiiess and chord of the 540 r o t o r blade were increased s l i g h t l y t o accommodate many blade-surface pressure transducers. It i s hoped t h a t f u t u r e i n v e s t i g a t i o n s of r o t o r s c a l i n g will make use of these f u l l - s c a l e simultaneom blade-surface and f a r - f i e l d acoustic data.
A photograph of the two-bladed, t e e t e r i n g , 1 / 7 geometrically scaled AH-l/OLS model r o t o r i s shown i n Fig. 4. The r o t o r was i n s t r u - mented with 50 miniature pressure transducers: 32 absolute flush- mounted K u l i t e transducers on one of t h e blades and 18 d i f f e r e n t i a l - pressure transducers on the second blade. The absolute transducer l o c a t i o n s were chosen t o match some of the r a d i a l and chordwise trans- The geometric ducer p o s i t i o n s i n the f u l l - s c a l e tests of Ref. [14].
c h a r a c t e r i s t i c s of the 1.916-m-diam (6.3-ft) model-scale OLS blades a r e shown i n Fig. 5 ( a ) , and the absolute transducer l o c a t i o n s are shown In Fig. 5(b). Some high-advancing-tip Mach number absolute 17-8
ORIGINAL PAdE 18
0 UPPER SURFACE
OF POOR QUALITY
e LOWER SURFACE
P
I I
I
,
i
L
1.
z e E L ? $ 5 % I R 6 1 PERCENT RADIUS (b) Scale model AH-1/OLS rotor-blade absolute pressure transducer l o c a t i o n s .
Fig. 5. Concluded.
blade pressure data are discussed i n t h i s paper. The r o t o r model i s d i r e c t l y scaled from t h e f u l l - s c a l e OLS blades. A s i n t h e f u l l - s c a l e OLS blade, model-blade thickness and chord have been increased s l i g h t l y over those of the standard BHT 540 AH-1 blade. Although t h e geometrical d i f f e r e n c e s between t b e 540 and t h e OLS blades are small, they have been accounted f o r i n t h e comparisons of t h e model OLS blade and the f u l l - s c a l e 540 blade acoustic measurements shown i n t h i s paper.
4. Acoustic Scaling Parameters Acoustic scaling, l i k e most procedures t h a t are dependent: on f l u i d dynamic processes, is governed by t h e fundamental laws of mass and momenrum. When t h e governing equations are placed i n nondimen- s i o n a l form and the important nondimensional parameters of noise r a d i a t i o n a r e matched, i t is possible t o duplicate a large-scale a c o u s t i c event i n small scale. A concise mathematical form (an i n t e g r a l equation) f o r the sound generated by bodies i n a r b i t r a r y motion i s given i n Ref. [15]. I n ihe Appendix of t h i s paper, t h i s equation is rewritten i n nondimensional form (Eq. (A?)) and t h e r e s u l t s i n t e r p r e t e d f o r model- dnd f u l l - s c a l e r o t o r t e s t i n g .
One of t h e f i r s t conditions f o r model-to-full-scale a c o u s t i c scaling i s geometric s i m i l a r i t y . Thus, a l l model dimcnsions are l / y t b e s t\e f u l l - s c a l e dimensions of the noLse generating surfaces: f u l l - s c a l e length = model-scale length 17-10 (A2) is that t h e model and f u l l - A second condition s p e c i f i e d by Eq.
scale hover t i p Mach ,iumbers (MH) be i d e n t i c a l . n u s , a geometrical y must be o f f s e t by m- raduction t o model-scale by the scale f a c t o r i n c r e a s e of rotor-shaft r o t a t i o n a l speed. Speed-of-sound d i f f e r e n c e s between model and f u l l - s c a l e conditions must be accounted for. There- fore, t h e model-rotor s h a f t r o t a t i o n a l speed & (subscript m f o r model) I s r e l a t e d t o t h e f u l l - s c a l e s h a f t r o t a t i o n a l speed by Eq. (A3): Also, a u n i t of model-scale time is r e l a t e d t o r ~ l l - s c a l e time by Eq. ( A 4 ) : 1 a0 A l l temporal d a t a shown i n t h i s paper have been normalized by a r o t o r revolution or f r a c t i o n thereof, thereby accounting f o r time-scaling d i f f e r e n c e s .
As discussed in the Appendix, f o u r n o n d h n s i o n a l parameters kave t o be matched i f model scale is t o d u p l i c a t e *he f u l l - s c a l e phe-
nomenon of i n t e r e s t . II? a d d i t i o n t o hover t i p Mach number %) whi.ch
was discussed above, they are advance r a t i o p , + h r u s t c o e f f i c i e n t CT, and tip-path-plane angle q p p . Of course, by matching f u l l - scale hover t i p Mach number and advt.we r a t i o , t h e advancing-tip Mach number (MAT) is automatically duplicated. For high-speed impulsive noise, advancing-tip Mach number i n ? t e a d of advance r a t i o is o f t e n chosen as the dominant nondhensiona: parameter. It c o r t r o l s , f o r the most p a r t , how a c o u s t i c waves are r a d i a t e d t o the I --field (Ref. [4]).
I f a l l of the governing nondimensional parameters of high- speed impulsive noise are matched, Eq. (A2) s t a t e s t h a t t h e a c o u s t i c pressure coeff, ' e n t [Ci(z,i)J of the f u l l - s c a l e r c t o r is t h e same as the a c o u s t i c pressure c o e f f i c i e n t of t h e model scale r o t o r fCim(Z,,&)] ; t h a t is, I f , a s is normally t h e case, it is d e s i r e a t o compare dimensional pressura- time-histories from two separate tests, where a l l t h e gov- erning nondimensional parameters have been matched, it is necess.nry t o a d j u s t the pressure l e v e l s t o account f o r d i f f e r e n c e s i n p,ai, which is proportional t o the ambient pressure Po, For convenience p a s t vind-tunnel a c o u s t i c p r a c t i c e s , a l l a c o u s t i c and i n concert with d a t a presented i n t h i s report w i l l be r e f e r r e d t o ISA standard day,
sea-level pressure (indicated by *). Thus, model-scale pressured
measured during wind-tunnel t e s t i n g become or, fn terms of aabient p r e s m r e r a t i o , Because t h e DNW is located near sea level and n o & operatfng condi- tivns were always close to t h e I S A standard &y, only d l pressure r a t i o c o r r e c t i o n s P /Po* were necsssary f o r t h e model a c o u s t i c data %I presentad i n t h i s paper.
Similar correction procedures are necessary for f l i g h t - t e s t Again. a l l acoustic data are r e f e r r e d to ISA standard day, data.
-&-level pressure by
P ' ( ' ; , E , P ' ( E , : )
(3)
P'*(it,t) = * *2'=
P0/P$ Poa:lPoso Because a c o u s t i c daca were gathered a t v-irying prassure a l t i t u d e s , c o r r e c t i w s were not i n s i g n i f i c a n t f o r f l i g h t test data ( P o / P : of t h e order 0.7 t o 0 . 8 ) .
Tnis was r e a l l y the f i r s t t i m e that such procedures had been applied to r o t o r acoustic testing. Previously it was thought (Refs. [4] and 16-11]) that s*mly a d j u s t i n g the measured pressures f o r density r a t i o (i.e., po/po) was adequate t o c o r r e c t pressures t o sea-level s t a n d x d conditions. However, t h e forraal transformation of the governing equation. to nondimensional form (presented i n t h e Appendix) has shown t h i s t o be only p a r t i a l l y correct. The complete adjustment r e q u i r e s c o r r e c t i o n s according to t h e bmbient pressure
r a t i o o r equivalent t o density - and temperatcre ratio (since
2 * * 2 , P o ~ l P o a o P O T O / P $ 3 The model-rotor test program i n t h e D N W comprised 3 wk of tunnel occupancy. Because t h i s was the f i r s t r o t o r acousti: test i n the free-jet of t h e DNU, 2 of the 3 wk were used f o r setup and in-place preoperational checkcut. This c a l l b r a t i o n and checkout phase consisted of d e t s c t i o n and attenuation of a c o u s t i c impulse r e f l e c t i o n s from the Eicrophone support s t r u t s , background noise measurements, and a high-speed hovering check with untwisted UEI-18 blades. This last check allowed the high-speed d a t a taken In t h f e f a c i l i t y to be campared with data Erom o t h e r a c o u s t i c f a c i l i t i e s throughout t h e world ( e - e . , Ref. [16]). Some r e s u l t s and d e t a i l s of these c a l i b r a t i o n procedures a r e presented i n Appendix B of Ref. [12].
I n general, the Dhi w a s shown t o have some of the b e s t r o t o r a c o u s t i c measurement c a p a b i l i t i e s i n t h e world. I n t h e remaining week of tunnel occupancy, a t o t a l of about 150 data p o i n t s were taken, cover- ing both blade-vortex-tcterac t i o n and high-speed impulsive noise operating conditions. The range of t h e four governing nondimensional 17-12
parameters (hover t i p Mach number h, advance r a t i o IA, t h r u s t
c o e f f i c i e n t C;r. and tip-path-plane acgle 3pp) covered were
% (0.55-0.72)
M M (0.73-0.94) p (0.1-0.35)
m N A L PAGE ?3
5 (0.0047-0.0065)
OF P O O R QuALm
aTpp (-5"-t7") The nondimensional t e s t i n g envelope of model-scale test conditions is shown i n Fig. 6 . A l s o presented in Fig. 6 are the dimensional rate- of-climb p r o f i l e s of a simple performance model f o r t h e AH-1 h e l i - copter (Cobra). The model-test d a t a presented i n t h i s paper emphasize high-speed impulsive noise and thus encompass many test p o i n t s in t h e r i g h t half of t h e cross-hatched region. For two reasons, many of * these p o i n t s were chosen to be r e p r e s e a t a t i v e of descending f l i g h t : (1) t h e f u l l - s c a l e data are, a t times, gathered in descending f l i g h t because of power l i m i t a t i o n s of t h e YO-3A a i r c r a f t ; and (2) a model- s h a f t - t i l t i a g c a p a b i l i t y was not a v a i l a b l e during the t e s t , restrict- ing t h e forward tip-path-plane ( T ~ ~ ) angles a t high forward speeds.
AOVANCE RATIO, , . t Fig. 6. OLS model t e s t i n g envelope.
17-13 5 . Mode l-Scale/Full-Scale Acoustic Comparisons The simple and most d i r e c t method of comparing model-scale and f u l l - s c a l e a c o u s t i c data is through a d e t a i l e d a n a l y s i s of acous- Such a comparison of AH-l/OLS model t i c pressure time-histories.
acoustic data with AH-1 f u l l - s c a l e data measured with the i n - f l i g h t method developed and performed by Boxwell and Schmitz (Refs. [ 9 ]
and [lo]) is i l l u s t r a t e d i n Figs. 7 and 8. Negative peak pressure
levels are p l o t t e d as a function of advancing-tip Mach number, t h e most important nondhensional parameter of high-speed impulsive noise.
The sketches i n Fig. 7(b) i l l u s t r a t e t h e equivalence of wind-tunnel ana i n - f l i g h t a c o u s t i c t e s t i n g and i n d i c a t e t h a t t h e d a t a shown were gathered by an in-plane microphone located 1.8 r o t o r diameters ahead of t h e hub. Because t h i s is one of t h e easiest station-keeping posi- t i o n s f o r t h e i n - f l i g h t data gathering method, the f u l l - s c a l e data shown are thought to be of high q u a l i t y . A l l of the individual high- speed in-plane d a t a p o i n t s are shown i n Fig. 7(b) (small circles and d o t s ) , i l l u s t r a t i n g a r e p r e s e n t a t i v e scatter band f o r t h e f u l l - s c a l e technique. The d a t a indicated by c i r c l e s were c o r r e l a t e d with advanclng-tip Mach number from t h e more unsteady runs. For these cases, t h e p i l o t made s u b s t a n t i a l adjustments t o t h e c o l l e c t i v e con- t r o l t o hold c u r r e n t r e l a t i v e s p a t i a l p o s i t i o n between the micro- phones mounted on t h e fixed-wing a i r c r a f t and t h e h e l i c o p t e r . These c o l l e c t i v e changes a l t e r e d t h e r o t o r rpm, thereby changing t h e instan- taneous advancing-tip Mach number. These e f f e c t s were accounted f o r i n the data shown by c o r r e l a t i n g the instantaneous a c o u s t i c p r e s s u r e s with instantaneous advancing-tip Mach number. The d o t s represent t h e ease of comparison, they are more steady f u l l - s c a l e d a t a points. For generally ir-dicated by the shaded area. The model-scale wind-tunnel d a t a are shown by the c i r c l e d crosses. I n Fig. 7, t h e model a c o u s t i c data have been adjusted f o r t h e small d i f f e r e n c e s i n noise path lengths (using a l / r decay l a w which is discussed later) and blade thickness. The f u l l - s c a l e data have been corrected t o ISA standard Eq. (3) i n the previous s e c t i o n day sea-level pressure according t o of t h i s paper.
The mode 1- scale / f u l l - s c a l e comparison of in-p lane microphone data is e x c e l l e n t . Peak negative p r e s s u r e s and waveform shapes agree remarkably w e l l over the e n t i r e Mach-number range t e s t e d .
Some s m a l l amplitude discrepancies are noted at t h e higher advancing-tip Mach numbers i f the model s c a l e data are compared with only the more steady f u l l - s c a l e data. Although t h e r e is no conclu- s i v e evidence t o account f o r these small discrepancies, the d i f f e r - ence between f u l l - s c a l e and model-scale Slade/hub dynamic character- i s t i c s is thought t o p l a y a r o l e . I n p a r t i c u l a r , some preliminary c a l c u l a t i o n s , performed by J. Corrigan of Bell Helicopter, i n d i c a t e t h a t simple lead-lag motion d i f f e r e n c e s r e s u l t i n g from trim changes could account f o r f u l l - s c a l e advancing-tip Mach number reductions of 0.006. Although reductions i n advancing-tip Mach number of 0.006 would be barely noticeable a t low forward v e l o c i t i e s , the e f f e c t could e a s i l y explain the small d i f f e r e n c e s between model- and f u l l - s c a l e acoustic data shown a t the high-advancing-tip Mach numbers.
17-14
oR(QIIyAL PAGE I S
01s MODEL ROTOR ACOUSTIC SlGNATURES
+------
I ' ,! [i
I
-100 -200 -300 -I+ . - > - - / - A ' b t
: o - - ' ; ' 1 7 , ! I
-200 -3CO VT, knots = 72 V I , knots = 96 VT, knots = 118 VT, knots * 143 VT, k M 8 e 150 RID. ft/min * 0 l/D, ftlmin= 0 R/D.ftlmm = 0 RtD.ft;min= 400 RtD.ft/mtn = 1000 p = 0.169 p - 0 2 2 2
p - 0.278 L d = 0.330
/I 0.348
CT 0.0054 CT 0.0064 CT = 0 . m CT - 0.0064 CT -0.0063
MAT = 0 801 MAT = 0.802 M ~ ~ = 0 7 6 9 MAT 0 878 MAT 9 0.898 I
L
(Cl A H 1s 1540 ROTOR) ACOUSTIC SIGNATURES Comparison of model and full-scale acoustic pressure for an Fig. 7 .
in-plane microphone 1.8 rotor diameters ahead. (a) Model-scale results for one rotor revolution; (b) wind-tunnel and in-flight acoustic testing equivalence; (c) full-scale pressure-time histories for one rotor revolution.
1.7-1 5 The importance of this e f f e c t w a s a l s o demonstrated during mo9-i t e s t i n g A n t h e DNW. After running a t a condition of high- advancing-tip Mach numbers and high thrust, the feathering bearings became worn, allowing additional lead-lag blade motions of 20.5'.
Acoustic data gathered under these conditions were reduced i n ampli- tude. by an average of 20% from those of a similar run taken with a gool? feathering bearing. It would appear t h a t in-plane dynamics can p L 7 a s i g n i f i c a n t r o l e i n t h e c o r r e l a t i o n between model and f u l l - 8c:Ile data. Fortunately, in the test reported here, the model and f u l l - s c a l e blades are s t i f f and heavy, thus minimizing these e f f e c t s .
However, f o r more modern r o t o r s , which are more l i k e l y t o be s o f t e r and l i g h t e r than the 540 r o t o r , the f i r s t mode of Lead-lag, torsion, and flapping should probably bo dynamically scaled.
The dramatic increase in peak pressure levels and the waveform sbr: )e changes t h a t occur near t h e delocalization advanchg-tip Mach aw:.)er are i l l u s t r a t e d i n t h e averaged time-history inserts of Fig. 7; F i t . 7(a) shows t h e model-scale results and Fig. 7(c) t h e full-scale pressure-time h i s t o r i e s f o r one r o t o r revolution. A d e f i n i t e change in pulse shape occurs in both the model-scale and f u l l - s c a l e data i n the 0.86 t o 0.9 advancing-tip Mach number range.
These changing waveform c h a r a c t e r i s t i c s are more c l e a r l y indi- cated in Fig. 3 where the time-scale oE one impulse has been length- A t advancing-tip Mach numbers of 0.864 and below, t h e high- ened.
speed impulsive noise waveform is almost symmetrical on both the A t s l i g h t l y higher advancing-tip nodel-scale and full-scale data.
Yach numbers (Mm : 0.885), t h e waveform of the model-scale and f u l l -
x a l e data begin t o change, becoming more saw-toothed. A t advancing- i p Mach numbers near 0.9, the waveform has changed t o a pronounced *.aw--tnothed shape; t h a t is, a large negative pressure peak is followed by a fiteep r i F . i n pressure (shock). I n e f f e c t , the l o c a l transonic f i e l d s of both t h e model-scale and full-scale have delocalized. A strong discontinuous pressure jump (shock) r a d i a t e s uninterrupted from t h e unsteady transonic flow f i e l d surrounding each rotor-blade
t i p to the acoustic f a r - f i e l d (Refs. (11, [ 4 ] , and [ 8 ] ) . The excel-
l e n t c o r r e l a t i o n i n waveform character i n the delocalization advancing- t i p Mach number region is a f u r t h e r demo.istration of t h e excellent scaling between modtl-scale and full-scale high-speed impulsive noise data.
It SI - Ald be noted t h a t the full-scale acoustic data presented i n t h i s p I e r have been averaged t o improve the signal/noise l e v e l and mor? completely define the character of the full-scale high-speed noise waveform, This was accomplished by signal-analysis techniques thz'. u t i l i z e the dominant negative-pressure peaks t o control the 6 ' g i t i t i n g process instead of u t i l i z i n g the l l r e v s i g n a l s from the ,lico,)ter. Becau3e the t a i l - r o t o r r o t a t i o n a l frequency is a non- integer m u l t i r , l ~ of t h a t of t h e main rotor. most t a i l - r o t o r noise is averaged out, leaving only the main-rotor impulsive noise signature f o r comparrson purposes.
Some differences i n the model-scale t o f u l l - s c a l e pulse shapeu c k e x i s t ; they occur s l i g h t l y before the l a r g e negative high-speed impulsive noise impulse, as i l l u s t r a t e d i n Fig. 8 .
These smaller 17-16 ORIG4NAl PAGE 1 3 OF POOR QUALITY L 1 . 8 D - d FULL SCALE ADVANCING TIP MACH NUMBER. M~~
/
!
/
O J J. PL-
\
I t
I id
-100 -
-200 - -300
MAT - 0.864
F i g . 8 . Comparison of model and full-scale acoustic waveforms for an in-plane microphone 1.8 rotor diameters ahead.
17-1 7 OHIGIIUAL PAGi S OF POOR QUACt'"Y differences i n waveform shape are known t o be strongly influenced by
blade-vortex interactions (Ref. I l l ] ) . Small changes or differences
between model- and full-scale tip-path-plane angles are thought t o be responsible f o r t h e waveform differences shown. A more c a r e f u l look into these e f f e c t s is planned.
The peak pressure level versus advancing-tip Mach number of two microphones is shown i n Fig. 9, using t h e right-hand vertical hl E .
z
-
-500 Ly a
d
z
-400 3
-
a w >, a -300
d
u) k a w
-200 5
w
E
s
-100 2
Q .77 .80 .85 .93 ADVANCING-TIP MACH NUMBER, MAT Fig. 9. Peak pressure decay rate f o r various operational conditions of the model rotor.
scale. The microphones a r e located along an imaginary l i n e t o t h e rotor hub d i r e c t l y ahead of the model r o t o r , t h e second microphone being exactly twice t h e distance from the hub as the f i r s t . The measured peak pressure r a t i o , Pi/Pi, is a l s o p l o t t e d (using t h e l e f t - hand scale) indicating a simple l / r decay law (spherical spreading) which i s independent of advancing-tip Mach number. This f i g u r e shows that in-plane microphone locat,ons greater than 3 r a d i i from the r o t o r hub a r e i n the high-speed impulsive noise acoustic f a r - f i e l d of t h e rotor. This r e s u l t has been used i n Figs. 7 and 8 t o correct t h e model-rotor amplitudes f o r minor measurement-position differences between model-scale and full-scale r o t o r data. The r e s u l t is a l s o used i n Fig. 10 t o correct full-scale data t o model-scale r o t o r non- dimensional distances. I n some full-scale positions, these correc- tions were not insignificant.
17-18 D i r e c t i v i t y comparisons of model-scale and f u l l - s c a l e data are presented i n t h e d i r e c t i v i t y p r o f i l e s of high-speed impulsive noise shown i n Fig. 10. The model-scale data shown i n t h i s f i g u r e were gathered under c o n t r o l l e d conditions and can be considered t o repre- s e n t q u a n t i t a t i v e d i r e c t i v i t y p r o f i l e s of r a d i a t e d high-speed impul- sive noise. However, the very n a t u r e of the i n - f l i g h t , f u l l - s c a l e data-gathering technique makes s i m i l a r measurements at o t h e r than a f e w c a r e f u l l y c o n t r o l l e d microphone p o s i t i o n s d i f f i c u l t a t best.
Measurement-position e r r o r s (azimuthal, elevation, and distance) and varying i n - f l i g h t operating conditions tend t o l e s s e n t h e accuracy of t h e results. However, i n s p i t e of these q u a l i f i c a t i o n s , t h e compari- sons i l l u s t r a t e d in Fig. 10 are q u i t e good, demonstrating t h e scala- b i l i t y of t h e high-speed impulsive noise.
Figure 10 p r e s e n t s the d i r e c t i v i t y of high-speed impulsive noise a t an advancing-tip Mach number of 0.84 (below d e l o c a l i z a t i o n ) .
The l o n g i t u d i n a l d i r e c t i v i t y , shown i n Fig. 10(a) is highly direc- as previously reported (Refs. [4], (81, and [lo]). The peak t i o n a l , negative amplitude of t h e nearly symmetrical pulse decreases r a p i d l y
a t increasing l o n g i t u d i n a l d i r e c t i v i t y angles (4) . The model-scale/
f u l l - s c a l e comparisons suggest t h a t the f u l l - s c a l e data shown at an estimated 30" below the h o r i z o n t a l might have been a c t u a l l y taken a t s l i g h t l y l a r g e r angles. Also shown i n t h i s f i g u r e is t h e microphone No. 1 waveform of t h e model-scale d a t a taken a t twice t h e d i s t a n c e of microphone No. 2. Although n e a r l y i d e n t i c a l i n shape, it is approxi- mately h a l f t h e amplitude of t h e c l o s e r microphone, as would be pre- d i c t e d from the l/r sound decay l a w .
A lateral d i r e c t i v i t y comparison a t t h i s same advancing-tip Mach number is shown i n Fig. 10(b). A t the in-plane microphone posi- t i o n s , the model-scale/full-scale comparisons are e x c e l l e n t . A t 30" under t h e r o t o r plane, the comparisons are only s l i g h t l y degraded.
Some of these waveform d i f f e r e n c e s are a t t r i b u t a b l e t o t h e l i m i t e d This f i g u r e azimuthal l o c a t i o n s where f u l l - s c a l e d a t a are a v a i l a b l e .
of the high-speed impulsive noise. A t confirms the focused nature t h i s advancing-tip Mach number (MAT = 0.84) and advance r a t i o ( p = 0 . 2 6 ) the maximum noise i n t e n s i t y is d i r e c t e d forward b u t t o the advancing s i d e of t h e r o t o r . A similar l o n g i t u d i n a l and lateral d i r e c t i v i t y comparison of high-speed impulsive noise a t a n advancing- (above d e l o c a l i z a t i o n ) t i p Mach number of 0.895 is given i n Ref. [12].
As i n the lower Mach number case presented here, t h e comparison !-?tween model-scale and f u l l - s c a l e is e x c e l l e n t .
Taken together, Figs. 7-10 conclusively demonstrate t h a t care- f u l l y designed and nondimensionally t e s t e d small-scale models can duplicate the high-speed impulsive noise generated by f u l l - s c a l e r o t o r s . As demonstrated, the e x c e l l e n t aerodynamic and a c o u s t i c t..vironment of t h e DNW makes i t more than adeqvate f o r high-speed impulsive noise t e s t i n g .
One oi the major advantages of wind-tunnel t e s t i n g over f l i g h t t e s t i n g is t h a t i t makes i t possible t o explore wide ranges of test conditions i n the relative s a f e t y and c o n t r o l l e d conditions of t h e wind tunnel. Figure 11 presents high-speed impulsive noise l e v e l s and pulse shapes f o r in-plane microphone No. 2 over a range of 17-19 0WG)NAL PAGE f8 O F POOR QUALITY
\
-50 Irl
w > L I II I / FULL SCALE MODEL SCALE MAT = 0.839+0.842 MAT 0.837
p - 0.270+0.276 p = 0.26
CT 0.0064 5 =o.w64
RID= 0 hlmin RID = 0 ft Imin V r * 116 knots VT= 116+118 knot$ (a) Longitudinal d i r e c t i v i t y (MAT = 0.837).
Fig. 10. Compar+.son of model and f u l l - s c a l e impulsive noise d i r e c t i v i t y .
17- 20 OR1QINAl PAGE 3 5
OF POOR QUALTPQ
IN-PLANE NOISE SIGNATURES FIS M/S
- I
-
-
-
!
e =oo
r v* = 00 /
\
,\
0 -29" $ 5 3 1 ' -100 30" DOWN NOISE SIGNATURES (b) Lateral directivity.
Fig. 10, Concluded.
17-21
ORIGINAL PAGE Is
OF POOR QUALm MH 0.698
MH - 0.887 MH 0.880
MAT 0.902 MAT * 0.831 MAT 0.887
W 0 7 5 . 3 mlr V = 76.8 mls W - 70.8 mlr
(= 147 knots) (= 163 knots) -1m N
I
a- W -10
t
u a
1 W
a & Y -1 ADVANCING-TIP MACH NUMBER, MAT Fig. 11. High-speed impulsive n o i s e v e r s u s advancing-tip Mach number a t constant advance r a t i o (model s c a l e ) .
advancing-tip Mach numbers from 0.73 t o 0.93. For a l l of the cases shown i n Fig. 11 the advance r a t i o p was held constant and t h e tunnel v e l o c i t y and r o t o r rpm were varied accordingly. Large changes i n peak pressure l e v e l and pulse shape a r e seen over a r a t h e r l i m i t e d range cf advancing-tip Mach numbers. As noted previously, delocaliza- t i o n occurs a t an advancing-tip Mach number of about 0.89. These new measurements i n the very favorable a c o u s t i c a l and aerodynamic environ- ment of the DNW confirm the f i z d i n g s of R e f . [ 4 ] , which show t h a t advancing-tip Nach number i s t b e key parameter of high-speed impul- s i v e noise. I n Fig. 11, the r a t e of increase of high-speed impulsive 17-22 noise l e v e l is p l o t t e d , using two scales: l i n e a r and logarithmic.
The l i n e a r s c a l e emphasizes t h e f a c t t h a t the r a d i a t e d noise increases Rowever, t o high l e v e l s with increasing advancing-tip Mach numbers.
t h e logarithmic p l o t i n d i c a t e s t h a t t h e rate of increase i n l e v e l does not continue t o increase. I n f a c t , the logarithmic p l o t reaches its l a r g e s t slope j u s t before t h e de! o c a l i z a t i o n advancing-tip Mach number (MAT 0.89) x d then begins t o f l a t t e n out. A similar result [ 1 6 ] f o r a hovering r o t o r , where it was found was reported i n Ref.
that the l o c a l transonic flow f i e l d tended t o weaken t h e rate of increase of acoustic l e v e l s a t zir above t h e hover d e l o c a l i z a t i o n Mach number. A similar mechanism is thought t o apply here. However, t h e unsteadiness of the transonic aerodynamic f i e l d w i l l undoubtedly influence the r e s u l t i n g acoustic r a d i a t i o n as w e l l . Variations of t h e peak pressure l e v e l s and waveforms with t h e other s c a l i n g param- eters (advance r a t i o , tip-path-plane angle, and t h r u s t c o e f f i c i e n t ) were investigated during the DNW t e s t i n g and are reported i n Ref. [12].
Their influence on t h e r a d i a t e d noise w a s confirmed t o be l e s s than t h a t of advancing-tip Mach number.
6.
Model-Scale Blade Pressures Throughout t h e DNW t e s t i n g , model rotor-blade pressure data were c o l l e c t e d simultaneously with the r o t o r a c o u s t i c s over t h e e n t i r e test envelope shown i n Fig. 6. As previously mentioned, t h i s envelope encompassed r o t o r operating conditions where both high-speed and blade-vortex i n t e r a c t i o n noise are known t o occur. On-line monitor- ing of a l l the blade-pressure data indicated t h a t it is of very high q u a l i t y over the e n t i r e matrix of f l i g h t conditions. I n keeping within the scope and purpose of t h i s paper, s e l e c t e d blade pressures f o r two high-speed test conditions are presented t h a t are of particu- lar i n t e r e s t a c o u s t i c a l l y . The two conditions a r e those f o r which acoustic signatures were shown previously i n Fig. 8. I n t h a t f i g u r e , model-rotor acoustic waveforms a t t i p Mach numbers of 0.864 and 0.896 were compared.
The 0.864 Mach number condition exhibited a nearly symmetrical waveform, whereas t h e 0.896 acoustic waveform was charac- t e r i z e d by a rapid (shock-like) pressure rise. The latter condition is s l i g h t l y above t h e d e l o c a l i z a t i o n Mach number and t h e former (0.864) is, s i g n i f i c a n t l y f o r acoustics, below it. Since t h e blade pressures and a c o u s t i c d a t a were acquired simultaneously i n t h i s test, t h e p o s s i b i l i t y of r e l a t i n g common c h a r a c t e r i s t i c s i n each was afforded. More s p e c i f i c a l l y , an i n t e r e s t i n g aspect of the delocali- zation hypotheris is t h a t t h e r e i s a high subsonic Mach number where shock waves t h a t e x i s t on t h e r o t o r blade escape t o t h e a c o u s t i c f a r - f i e l d . mhese waves a r e confined, however, t o a region surrounding the r o t o r t i p a t Mach numbers only s l i g h t l y below t h e d e l o c a l i z a t i o n Mach number but escape t o t h e acoustic f a r - f i e l d a t Mach numbers above it. I n both cases, shock wave3 e x i s t near t h e a i r f o i l surface and, theref ore, should be i d e n t i f i a b l e i n the measured blede-pressuro data.
Blade absolute pressures f o r the lower advancing-tip Mach number (0.864) a r e show.1 i n Fig. 1 2 f o r one r o t o r revolution. The pressures a r e measured on t h e upper surface near the blade t i p a t 95.5% radius, are referenced t o atmospheric pressure, and averaged ? 7-23 ORlGlNAl PAGE !S OF POOR Q U A L m
MA* - 0.864
p = 0.298 95.5% R UPPER SURFACE %CHORD
----- 15
- - 25
-.- 40
-..- &
-0- M)
- 70
I I I I 1 I I I 1 I I I 330 360 180 210 240 270 300 30 60 90 120 150 AZIMUTH ANGLE, deg Fig. 12.
Model OLS r o t o r upper-surface blade pressure v e r s u s azimuth for MAT = 0 . 8 6 4 , 95.5% R.
64 times, Pressure-time histories a t chordwise l o c a t i o n s from 15% t o 70% chord are shown i n t h e individual curves of t h e f i g u r e . S t a r t i n g a t 0' azimuth (blade pointing downstream) a l a r g e r a p i d decrease i n pressure occurs near t h e leading edge (15% chord) and moves rearward on the blade ( t o 25% chord) a6 the blade advances t o t h e 45" azimuth p o s i t i o n . Beyond t h i s point as t h e blade approacheu 90°, t h e r e is a l a r g e drop i n the negative pressure peak as indicated by t h e 35% chord transducer.
T h i s is c h a r a c t e r i s t i c of a shock formation between 25% and 35% chord a t the 90' position.
As t h e blade slows 17-24
-1.6 -
from the maximum $ = 30" lb =Qo"
-
advancing-tip Mach
-1.2 -
number, the pressure
-
discontinuity be tween
25% and 35% chord -.8 -
becomes stronger u n t i l CP
-
t h e blade reaches 120' -.4 -
*a where t h e r e 3.s a rapid ' @ a
B **
collapse i n pressure 0 -
-
as t h e shock moves 1 1 1 1 )
forward and p a s t the . 4 L ' ' ' '
-lXi l i s h e d between 25% and
[
w =60 35% chord a t 45' - -1.2 azimuth. A t go", the
- * *
s t r e n g t h has decreased
-.8 -
but grows again u n t i l cP the blade passes 120'.
Both Figs. 1 2 and 13 -.4 5 i n d i c a t e the formation of blade pressure , ; * , discon t inu i t ies near .1 the rotor-blade t i p .4 a t an advanclng-blade- -1.6- t i p Mach number of I J J = 76' J, = 135'
F
0.864. Since the f a r -
- L
-1.2 f i e l d a c o u s t i c waveform
-
(Fig. 8 ) is nearly
-
-.8 i t would symmetrical, CP appear t h a t t h i s - a
-.4 - 0
discontinuous blade- *e
pressure disturbance **
0 1 **
is not preserved i n the r a d i a t e d acoustic t l k l ~ l .4 0 .2 .4 .6 .8 1.0 signature.
0 .2 .4 .0 .8 1.0 xlc xlc I n a s i m i l a r manner, the upper- Fig. 13. Model OLS r o t o r upper-surface surface blade-pressure chordwise pressure c o e f f i c i e n t versus d i s t r i b u t i o n a t the azimuth f o r MAT - 0.864, 95.5% R.
17-25 higher advmcdng-blade-tip Hach number (0.8%) is presented in Flgs. 14 and 15. Once again a large negative-pressur6 region is formed on the blade upper surface and emves tearvard as the rotor blade LpprOache8 the 90" position. A t this higher Ha& number, ha+ ever, the supersonic flow region extends farther back an the blade.
Figure 14 abows that a large pressure discontinuity (rise) now occurs between 40% and 452 chord at the 90" position.
A s before, both Figs. 14 and 15 show that the shock remains on the airfoil during MAT = 0.8s p = 0 . 3 4 5 95.5XR UPPER SURFACE % CHORD
-...--- 15
--25
....... 0 . 35
-*- 40
-..- g j
-0- 50
- 70
1 I I I I I I 1 I 1 I I 3u 60 90 120 150 180 210 240 270 300 330 380 AZIMUT); aNGLE, deg Fig. 14. Model OLS rotor upper-surface I * \s-i, z versus azlmuth fcr MAT = l.896 17-26 deceleration p a s t r -1.6 6 =30" J/ =90°L) 90" u n t i l a t 125"
the shock h a s moved -
forward again and
-
collapsed t h e l a r g e
cp --8 - *e
negative-pressure e
-1-21
regiou. It should e be noted t h a t at t h i s high-f orward-speed e
--L-
condition, t h e t i p - path-plane tilt w a s ' l e l ' 1 .4 limited (as explained in an earlier sec- t i o n ) so t h a t the r o t o r is i n a descent of 800 ft/min. The i r r e g u l a r i t i e s i n t h e pressure d i s t r i - butions of Fig. 14 i n the v i c i n i t y of 60" azimuth are believed t o be t h e r e s u l t of blade/wake
i n t e r a c t i o n s . r
- 1 . 6 ~ =6(J ; . '20 Unlike the
-
-1.2 -
lower Mach number e e case (0.864) i n
-
-.8 -
e which strong dis- - e
-
cP e .
c o n t i n u i t i e s e x i s t e d -.4 8 on t h e blade b u t not *e e i n t h e radiated e *e 0 - 8 acoustic signature,
- -
the advancing-tip , 1 1 1
.4 -
Mach number case of -1.6- 0.896 (Fig. 8 ) 6 = 1 3 5 i j =75
-
e x h i b i t s shock -1 i k e
-
-1.2 disturbances both on the blade and i n e * the f ar-f i e l d acous- - . 8 - CP t i c waveform, indi-
-
c a t i n 8 t h a t above t h e -.4
l e e
*e d e l o c a l i z a t i o n Mach e number the l o c a l 0 : shocks t h a t e x i s t
1 1 1 1 J 1 I e ~ e e l e l I
ne -r the rotor-blade .4 surface r a d i a t e as shock waves t o the acoustic f a r - f i e l d .
17-27 7. Concluding Remarks Acoustic and blade-pressure data taken i n the world's l a r g e s t
anechoic wind tunnel - the DNW i n the Netherlands - have documented
the high-speed noise r a d i a t e d from a 1/7-scale model main r o t o r of the AH-1 series h e l i c o p t e r . The d a t a confirm and expand many of t h e The major f i n d i n g s are known f e a t u r e s of high-speed impulsive noise.
as follovs: A set of nundimensional s c a l i n g equations w a s developed 1) These equations from the governing equations of r o t o r acoustics.
r o t o r a c o u s t i c were successfully used t o compare small-scale model data taken a t sea level i n wind tunnels with f u l l - s c a l e a c o u s t i c d a t a measured i n - f l i g h t a t a l t i t u d e .
High-speed impulsive n o i s e model-scale amplitudes and 2) waveforms compare exceedingly w e l l with f u l l - s c a l e i n - f l i g h t a c o u s t i c a wide range of advancing-tip Mach numbers and d i r e c t i v i t y d a t a over angles. These r e s u l t s conclusively demonstrate t h a t model-scale r o t o r s can be used t o explore p o t e n t i a l a c o u s t i c design chax'ges on f u l l - s c a l e helicopters.
High-speed impulsive noise is a highly d i r e c t i o n a l phe- 3)
nomenon - e n e r m is r a d i a t e d predominantly in-plane i n t h e d i r e c t i o n
of forward f l i g h t toward the advancing-blade s i d e of the r o t o r .
Advancing-tip Mach number is the dominant no3dimensional parameter t h a t governs high-speed impulsive noise r a d i a t i o n .
Blade pressures show t h a t l o c a l shock waves e x i s t near t h e 4) t i p of the r o t o r blade and "delocalize" t o t h e a c o u s t i c f a r - f i e l d above a "delocalization" hdvancing-tip Mach number. The mechanisms known t o c o n t r o l d e l o c a l i z a t i o n f o r the hovering transonic r o t o r a l s o appear t o s i g n i f i c a n t l y influence t h e forward-€light transonic acous- t i c radiation. Unsteady transonic e f f e c t s appear t o exert a secondary influence on the a c o u s t i c r a d i a t i o n when compared with t h e e f f e c t of advancing-tip Mach number.
Microphones located a t a distance of 3 r a d i i from the r o t o r 5 ) hub are i n the acoustic f a r - f i e l d of high-speed impulsive noise. A t g r e a t e r distances, the peak negative pressure l e v e l decays according t o a l / r l a w over the range of advancing-tip Mach numbers t e s t e d (MAT = 0.7-0.94).
For high-speed impulsive noise, signal-analysis techniques 6 ) can be used t o improve the signal-to-noise r a t i o of thm i n - f l i g h t - data.
By synchronizing with the l a r g e f e a t u r e s of the lmpulsive waveform, i n - f l i g h t a c o u s t i c d a t a can be averaged. The r e s u l t i n g waveform does not contain bothersome t a i l - r o t o r p e r i o d i c noise; i t r -$resents the amplitude and waveform from the main r o t o r only.
Perhaps the most important contribution of the present e f f o r t is the c a r e f u l documentation of the high-speed impulsive noise wave- forms. The q u i e t ambient and nearly anechoic p r o p e r t i e s of the DNW have minimized acoustic d i s t o r t i o n s . It is hoped t h a t the high qual- i t y of the r e s u l t i n g data w i l l be used t o guide the development of 17-28 theory. The data can a l s o be used t o help i n t e r p r e t s i m i l a r a c o u s t i c d a t a under less i d e a l conditions.
Ap2endi.x: Scaling of Acoustic Pressures The i n t e g r a l equation f o r the a c o u s t i c f i e l d generated by moving surfaces is given by (Ref. [15]) where and subscript o denotes ambient conditions. This equation contains t h r e e types of a c o u s t i c sources which are discussed i n d e t a i l i n Ref. [15]. The f i r s t i s a "monopole" source, which is governed by t h e time-rate-of-change of f l u i d mass displaced by t h e moving surface; j t i s known t o be a contributor t o high-speed impulsive noise. The second term i s a "dipole" source, which is dependent on a spatial d e r i v a t i v e of l o c a l s i r f a c e f o r c e s and is known t o be an important contributor t o blade-vortex i n t e r a c t i o n impulsive noise. The t h i r d term is a "quadrupole" source, which is governed by two s p a t i a l d e r i v a t i v e s of t h e Q i j stress tensor i n t h e volume of f l u i d sur- rounding the blade.
To put Eq. (Al) i n nondimensional form, define t h e following nondlmensional parameters: Nondimenslonal t i m e :
-
t = - (observer time) 2n /n T 7 s - (source t i m e or retarded time) 2nlSb Nondimensional geometry: Assuming t h a t v e l o c i t i e s normal t o the rotor-blade surface, vn, can be represented a s 17-29 where X = local surface slope, and remembering that
M = - '
3 Mach number of t h e flow over t h e blade a0
Defining % = r o t a t i o n a l t i p Mach number f m/aOs
Eo-ration (Al) becomes where
P m i m j + CpijM2 - - P 6 i j
'Qij = P, PO
and Equation (A2) d e f i n e s a nondimensional a c o t s t i c pressure c o e f f i c i e n t at a measurement point i n terms of nondlmensiona: parameters. Given unique values of a l l the nondimensional parameters on t h e right-hand si& of Eq. (A2), a unique value of $(a,;) is ensured. However, i t should be noted t h a t other governing nondimensional parameters are i m p l i c i t l y defined i n t h i s process.
Equation (A2) may be used t o develop d i r e c t l y s c a l i n g proce- dures and r u l e s f o r r o t o r t e s t i n g . Consider two d i f f e r e n t sized b u t geometrically similar r o t o r c of radius R, one f u l l scale and a seconds l/y scale. We s h a l l l e t R
y E scale f a c t o r = -
R, 17-30
ORlOlMl. PAGE IS
OF P W R QUALW where t h e s u b s c r i p t m denote8 model scale. An important nondlmen- s i o n a l parameter for acoustic s c a l i n g is r o t a t i o n a l t i p Mach r:w.ier, MT, To hold r o t a t i o n a l t i p Mach number t h e same for model and f u l l scale, the r o t a t i o n a l s h a f t rate must be adjusted so t h a t Because nondimensional time must a l s o be scaled, h = - - I n addition, Eq. (A2) r e q u i r e s t h a t M be scaled. If we consider t h e t i p c?f t h e r o t o r and neglect the spanwise flow along t h e blade, then (A s i m i l a r argument could be made a t any blade r a d i a l s t a t i o n . ) This implies t h a t t h e advance r a t i o p must be scaled, t h a t is, Thus, Equation (A2) a l s o r e q u i r e s t h a t be scaled f o r both Cpij and CQif model and f u l l scale. This implies s i m i l a r i t y i n the aerodynamic a t each flow f i e l d and s c a l i n g of r o t o r t h r u s t along t h e blade azimuthal angle. This requirement is approximated by maintaining similar in-flow through the rotor d i s k by means of s i m i l a r tip-path- plane angles and rotor-thrust c o e f f i c i e n t s .
The process of geometric scaling implies t h a t a l l l e n g t h s a r e scaled by y: 17-31 I n p r a c t i c a l terms, t h i s implies t h a t model measurement microphones should be positioned y times c l o s e r t o the hub c e n t e r than f u l l - scale geometric distances.
Acknowledment s I n a test of t h i s magnitude, t h e r e are many peaple who r i g h t - W e o f f e r our s i n c e r e s t f u l l y deserve recognition f o r a job w e l l done.
thanks t o our f r i e n d s and colleagues a t t h e Army Aeromechanics of Laboratory, N A S A Ames Research Center; t h e Acoustics Department t h e DFVLR, Braunschweig; the DNW Foundation and its parent organiza- t i o n s , DFVLR and NLR; NASA Ames Research Center; and the Army Struc- t u r e s Laboratory, NASA Langley Research Center.
References
F. H. Schmitz and Y. H. Yu, Transonic Rotor Noise - Theoretical
and Experimental Comparisons, Vertica, Vol. 5 , 1981, pp. 55-74.
M. P. Isom, Some Nonlinear Problems i n Transonic Helicopter Acoustics, Poly M / A E Report No. 79-19, Polytechnic I n s t i t u t e of New York, Biooklyn, N.Y., May 1979.
D . Hawkings, Noise Generation by Transonic Open Rotors, Research Paper No. 599, Westland Helicopters Limited, Yeovil, England, June 22, 1979.
F. H. Schmitz, D. A. Boxwell, and C. R. Vause, High Speed Heli- copter Impulsive Noise, Journal of t h e American Helicopter Society, Vol. 22, No. 4, O c t . 1977.
F. H . Schmitz and Y. H . Yu, Theoi-etical Modeling of High-speed Helicopter Impulsive Noise , Journal of the American Helicopter Society, Vol. 24, No. 1, 1979.
H. S t e r n f e l d , C. Bobo, D . Carmichael, T. Fukushima, and R. Spencer, An I n v e s t i g a t i o n of Noise Generation on a Hovering Rotor, P a r t 11, Report D210-10550-1, The Boeing Co., Vertol Div., Philadelphia, Penn., Nov. 1972.
H. S t e r n f e l d and E. Schaeffer, An I n v e s t i g a t i o n of Rotor Harmonic Noise by the Use of Small Scale Wind Tunnel Models, N A S A CR-166338, 1982.
F. H. Schmitz and D. A. Boxwell, In-Flight Far-Field Measurement of Helicopter Impulsive Noise, Journal of the American Helicopter Society, Vol. 21, No. 4, Oct. 1976.
D . A. Boxwell and F. H. Schmitz, Full-scale Measurements of Blade- Vortex I n t e r a c t i o n Noise, Journal of the American Heiicopter Society, Vol. 27, No. 4 , Oct. 1982.
17-32 D . A . Boxwell and F. €I. Schmitz, In-Flight Acoustic Comparison of 10)
the 540 and K747 Main Rotors f o r the A H - 1 s Helicopter, Appendix t o
U . S . Army Aviation Engineering F l i g h t Activity Report 77-38, Edwards AFB, Calif., Oct. 1979.
F. H . Schmitz, D. A. Boxwell, S. Lewy, and C. Dahan, A Note on t h e 11) General Scaling of Helicopter Blade-Vortex I n t e r a c t i o n Noise, Presented a t the 38th Annual National Forum of the American Helicopter Society, Anaheim, Calif., May 1982.
12) W. R. Splettstoesser, K . J. Schultz, F . 8. Schmitz, and D . A.
Boxwell, Model Rotor High-speed Impulsive Noise -Parametric Variations and F’ull-Scale Comparisons, Presented a t the 39th Annual National Forum of the American Helicopter Society, S t . Louis, Mo., May 1983.
J. C. A. Van Ditshulzen, G. D. Courage, and R. Ross, Acoustic 13) C a p a b i l i t i e s of the German-Dutch Wind Tunnel, DNW, Paper No. 9.5, Presented a t t h e 8 t h European Rotorcraft Forum, Ah-en-Provence, France, 1982.
G. A. Shockey, T. W. Williamson, and C. R. Cox, Helicopter Aero- 14) dynamics and Structural Loads Survey, Paper No. 1060, Presented a t the 32nd Annual National Forum of the American Helicopter Society, May 1976.
J. E . Ffowcs-Williams and D. L. Hawkins, Sound Generation by 15) Turbulence and Surfaces i n Arbitrary Motion, Philosophical Trans- a c t i o n s of the Royal Society of London, S e r i e s A, Vol. 264, 1969, pp. 321-342.
D. A. Boxwell, Y. H . Yu, and F. H. Schmitz, Hovering Impulsive 16) Noise: Some Measured and Calculated Results, Vertica, Vol. 3, NO. 1, 1979, pp. 35-45.
17-33