1980_nasa report_aeroacoustic wt tests of light aviation airplane with free or shrouded pusher...

8/11/2019 1980_NASA Report_aeroacoustic WT Tests of Light Aviation Airplane With Free or Shrouded Pusher Propeller

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Acr oacous ti c W ind -Tunnel T ests

of a L ight Twin- Boom G ener al-

Aviation Airplane With Free-or Shroud ed-Pushe r Propell ers

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: I Tes t s have been co ndu ct ed in tlle Langl e y Full-S c ale Tunnel to de t ermine_ the aerodynamic and a co u s t ic c hara ct eris t ics o f four different pusher - propeller

': arrangements o n a tw i n-boom, _ener a l -avia tio n air pl ane. The pr o peller sin c luded a 2-blade free propeller, two 3-blade shrouded propellers, and a5 -blade shrouded propeller•

,_ .... The tests were conducted for a range o f airplane angles of atta c k from 0°;_'_' to about ]6° for test speeds from 0 to about 36 m / sec (]]8 ft / sec) and for a._ r a nge of p£opeller-bl a de an g les and rotation speeds.

The free propeller provided the bes t overall aerodyna mic propuls i ve per-fo r mance. For forward- f light c onditions , the free-propeller noise levels werelower than those of the shrouded propellers, a nd in the static conditions , thefree-propeller noise levels were as low as those for the shrouded propellers,

except for the propeller in-plane noise where the sh-'ouded-propeller noiselevels were lower•

,,o'_ INTRODUCTION

• In recent years, increased e mphasis has been placed on the devel o pment° _ of propulsion systems which red u ce t he noise and engine-emissions poll u tion

of c ivil a irpl a nes without overly penalizing aerodynami c performance. Oneproposed method of accomplishing this objective for ligh t , propeller-driven,

i, general-aviation airplanes is the u se of a small-diameter shrouded prop_11er_,:'i with a direct-drive r o tary engine. Such a pr o pulsive system is expe c ted to

offer several advantages over the free-pr o peller system, including: (1) a_ :: more compact pr o pulsion package, (2) acoustic shielding, (3) minimization of

pollutants, a nd (4) less weigh t

_ The present investigation was conducted to determine the aeroacousti ci characteristics o f a preliminary design of a prop o sed shro u ded-propeller c o n-_<_'_' fig u ration whi c h wo u ld be u sed in s uc h a pr o p u lsion system. The tests were

conducted in the Langley Full-Scale Tunnel with a twin-boom, general-aviationo " airplane powered by three different shro u ded-p u sher-propeller configurations

and a fre e -p u sher-pr o peller c onfig u ra t i o n. Aero a co u stic dat a were o btainedfor airspeeds ranging from 0 to 36 m / se c (I]8 ft / sec) for a range of propeller-

: _ ..... blade angles and advance ratios.

..... i SYMBOLS AND ABBREVIATIONS

IAerodynamic data presented herein are referred to the stability system of

axes shown in figure I . The moment referen c e center was located horizontally_ at 37.5 percent of the mean aerodyn a mic chord and vertically at water line

1. 3 99 m (4.59 ft). Dimensional quantities are given in the Intern a tion a l

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Systc, m of. Units (S[), with U.S. Cu._ ; tomary Units g[w , n [n parenthese:_. Defini -t io ns and c o nversion [actors b t ,tween th e s y stems are presented i n refei : en c e 1.

B numb e r of pl : o_>l]or blades

' _ / Ill BW ban dw id th

C D d['ag coefficient, Drag / qS

C L lift coeff relent, Lift / qS

Pitching m o mentC m pitching-moment coefficient,

qS_

2 _QCp pr o peller p o wer co efficient,

pn2D 5

T

C T propeller thr u st coefficient,pn2D 4

TI

C T airplane thrust c oefficient ,qS

<t_ ? : c mean aerodynamic chord, 1.490 m (4.89 ft)

_-_... C t mean aerodynamic chord o f horiz o ntal t ail, 1.01 8 m ( 3 .34 ft )

D propeller diameter, m (ft)

d local propeller chord, m (ft)

d e propeller equ i valent ch o rd, -- dr 2 drR 3 •

FM propeller figure of merit; ratio ,f power required by an actuator

disk producing same thrust _ , power required by propeller:

0. 798CT3 / 2 0. 564CT 3 / 2for free pr o peller; .............. for shr ou ded propeller

Cp Cp

VJ pr o peller advance rati o , -- -

rid

M t propeller tip Mach number i

._: NRe Reynolds number

: 7 .',5_ <_:1

< _

-- 1980010746-T



In propeller revol u tions per second

Pb brake power, watts (ft-lb / sec)

P S shaft po wer , watts (ft - lb / se c)

p a c o u sti c pressure, N / m 2 (ib / ft 2)

P o re f eren c e a c oust ic pressure, 0 . 2 _N / m 2 (4.] 9 x 1 0 - 7 i b / ft 2 )

Q prop e ller t o rque, N - m (ib-ft)

q tunnel d y nami c pressure, N / m 2 _ib / ft 2)

R propeller r a dius, m (ft)

r radiu s st a tio n o f p r o pel le r , m (f t)

S wing area, ]6 258 m 2 (175.0 0 f t 2)

PS P L sound pressure level, 20 log -- , dB

P o

T effe c ti v e propeller thr u st , Dragprop off - Dragprop operating

t propeller-blade thickness, c m (in.)

V velo c ity , m/ sec (ft / sec)

W airplan e weight, kg

WL water l ine

wnmlg windmi l ling

X,Y , Z s t abilit y sy s t em of a xes

angle of at ta ck, deg (propeller shaft axis is used a s zero reference)

8 propeller-blade angle at any r a dius st a tion, deg

80 propeller-blade referen c e angle at 0.76 8 R (free propell e r) and0.803 R (shrouded propeller), deg

_f flap-deflection angle, deg

q pr o peller pr o pulsive efficiency , \Cp / \n_

0 mi c roph o ne a c ousti c angle , deg

3



P mass density of air, kg / m 3 (slugs / ft 3)

_: , _ :, _ micr o phone azimuth angle, deg

_ 2B 2-blade free propeller

I 3B 3-blade shrouded propeller, normal tip

3BT 3-blade shro u ded propeller, unloaded tip

5BT 5-blade shrouded propeller, unloaded tip

AI RPLANE AND TEST S ETUP

The airplane used in the tests was a modified version of a twin-engine,

twin-boom, general-aviation airplane. A three-view sketch of the airplane is

shown in figure 2; geome t ric charac t eristics of the free and shrouded propel-

lers are shown in figure 3; and photographs of t he airplane mounted for tests

in Langley Full-Scale Tunnel are presented in figure 4.

The forward propeller and engine of t he airplane were removed and the aft

{ fuselage contours were modified so as to be compatible with a rotary combustion

engine and a shrouded propeller. A direct-drive electric motor was used for

present tes t s to minimize engine noise and thereby permit measurements ofhe

the noise prod u ced by the propulsor unit. The airplane was designed to accept

either a free-propeller or a shrouded-propeller arrangement as shown in fig-

ure 3. In addi t ion to a conventional 2-blade free propeller, the tes t s

included two 3-blade shrouded propellers and a 5-blade shrouded propeller.

Some of the important geometric characteristics of the propellers were:

Diameter, Equivalent chord,:.. . Pr o peller m m Solidity

. .: , :::,:_ 2-blade free I .981 0.1 50 0. 098

.. (6.50) (0.50)

............3-blade shro u ded, uormal gip 1.006 0.140 0.266

(3.30) (0.46)

3-blade shrouded, unloade d tip 1.006 0.140 0.266

(3.30) (0.46)

........ '' I

5-blade shro u ded , unloaded tip ].006 0.091 0.289

One of the 3-blade propellers and the 5-blade prop el.let were twisted

abruptly near the tip (unloaded) whereas the other 3-blade propeiter had a

4



more c onventional (normal) twist distribution from root to tip. The reduced

tip loading was designed by the manufa_turer to alleviate acousti c al problemswithout unduly affecting the aerodynamic [_rforman e e. Blade-form curves forthe test propellers are presented in figure 5. The shrouded propellers hadNACA ]6-series airfoil sections , and the free propeller had Clark Y sections.The direction of propeller rotation was clockwise when viewed from the rearfor all of the test propellers.

The shrouded propellers originally had a diameter of 1.015 m (3.33 ft);however , during initial static tests , the propeller tips scraped the innershroud surface as a result of flexibility in the shroud and shroud supportsystems. The diameter of the shrouded propellers was therefore decreased by

" 0 .9 53 cm ( 0. 3 7 5 i n. ) t o i nsu r e ade q ua t e tip c learan c e for the remaining t es ts.Based upon data presented in reference 2, the tip c learan c e would be expe c tedto reduce propulsive thrust by about 1.5 per c ent.

P ressure- t ype microphones with nose cones were used in the a c ousti c por-

tion of t he investiga t ion. As shown in figure 4, II mi c rophones were used, andthe mi c rophones were mounted on 1.52-m (4.99 ft) stands placed in a semicircu-lar array around the right side of the airplane at a radius about the propelleraxis of 5.79] m (19.00 ft), as depi c ted in figure 6. The propeller axis wasabout 3.322 m (10.90 ft) above the m ic rophone array when the airplane was ate = 0° . As shown in figure 6, the acoustic angle 0 (the angle be ween the

hub-microphone line and t he propel l er axis) and t he azimuth angle _ weregenerally different for each microphone position. For the present tests, theairplane support struts were wrapped with 1.27-cm- (0.5 in.) thick wo o l-pilematting (fig. 4), and a special sound-absorbing ground board was used whi c hconsis t ed of a 10.16 - cm- (4 in.) thick layer of fiberglass c overed with40-percent porosity, 0.]58-cm- (0.0625 in.) thi c k perforated plate with0.3175-c m- (0.12 5 in.) di a meter holes (fig. 7).

CORRECTIO NS

The ae r od y nami c data have been c orre c ted for su p port strut tares , buo y -an c y , and airflow angularity. Wall c orre c ti o ns have been applied som e whatarbitrarily be c ause of the unique instal l ation of the airplane in the wind tun-nel. Normall y , tests simil a r to the p resent one are c ondu c ted with a nonporou sground board installed in the tunnel beneath the model and with wall c orre c tionsapplied to the data as obtained by the theor y of referen c e 3 . Sin c e the pres-ent tests were c on c erned with a c oustic as well as aerodynami c performan c e , the

ground board was c overed with acousti c devi c es as explained in the previoussection. Sin c e this a c ousti c material did not provide a finite , hard boundary

(su c h as the original ground board) nor was the tunnel jet c ompletely open (allfour walls unrestrained by solid boundaries) , the approa c h to applying meaning-

ful wall c orre c tions to the subje c t data was subje c t to considerable question.There was no c onvenient way to determine the wall c orrections experimentally ,so c al c ulati o ns were made to estimate the lift c hara c teristi c s of the basi c

unpo wered airplane , and these c alculated results were then c ompared with uncor-re c ted wind-tunnel test data and flight test data for an original Cessna 3 27

airplane. The wind-tunnel test data were then c orre c ted by methods of refer-ence 3 , using both the open-test-se c tion approach and the approa c h with one

5



s olid boundar y . The wall correction method that resulted in the best correla-tion of wind-tunnel and flight data a nd of calculations was that for an open-

te s t- s e c ti o n tunne l, s o the wall co rre c t io ns o f the s ub J _. c t rep o rt werp a ppli e das determined for the open test s e c tion witho u t a g round board.

TESTS AN D METHODS

Aer ody nam ic Te s t s

The a e r o dyn a mic tests wer e co nd uc te d for a s p e e d r a ng e f r o m 0 to a bo u t36 m/ s e c (1 1 8 f t / s e c), w ith m ost o f th e tests co n d u ct ed at a bou t 29 m/ se c

( 95 f t / s e c ) . Hi g h t unnel sp e e d s we r e fo und to b e unne c e ss ar y fr o m a R e y n oldsn u mb er sta n d poin t, a nd s t r uc t u r al l o a ding o f th e ai r f r ' -m e was min i mi z ed b y t her elatively low s p e e d tests . The l o we r spee d wa s a lso mette r fo r th e ac o u st i cmeasu r e me n ts s in c e t h e tu nn e l am b ie n t no i s e l eve l wa s co n si d e r a b ly l o we r f o rth e re d u c e d s p ee d .

In o r der t o c o n d uct the z e r o - s peed tests ( s tatic tests), a c l oth c urtai nwa s dr awn a c r os s the win d- t u n n el t e st- s e c ti o n e n tra n c e th r o a t t o p r e v ent th epro p e ll e r s lip st r e am fro m c i rc u lat in g th ro u g h the tu n n e l pass ages a nd th u sb a c k to the t e st p r op e lle r. The s tatic t es t s we re c o nd u cte d fo r a r a n g e ofpr o p e ll e r - b la de a n gles a nd r e v ol u t ion spee ds . T h e stat i c-t e st v a r ia bl e s f o rth e v a r io u s pr op e lle r a rr ang e men t s i n v e sti g at e d a r e g i v en i n table I .

Wi th t h e w in d tu an e l ope r at ing , tests we re co n d u ct e d for a r a ng e of a ngl e so f attac k f r o m 0° to 1 6 ° . P r opell e r- b la d e ang le an d r ev ol ution sp eed wer eva r ie d f r om a win dm illi n g con d ition to the r e v ol u tion s p ee d f or maxi mu m all o w-a ble mo to r c u rr ent, b u t n o t e xce e d i n g 270 0 r p m fo r the f r ee p r opelle r a nd5 00 0 rpm for t h e d u cte d pr o p e ll er s . I n ge n e ral , ma xi mu m mo t _) r c urr e n t r e s u lte di n a va lu e of po we r ve r y n e ar th e de sig n va l u e r e q u ir e d fo r the a ct u a l air p l a ne

(i . e . , 1 34 k W ( 1 80 hp) a t 2 7 0 0 rpm fo r the fre e p ro p e ll er o r 1 38 kW (]85 hp) a t5 00 0 r p m f or th e d u ct e d p r opelle r ) . Pr op e lle r dr i v e p o wer fo r the subjectte s ts wa s d etermined b y re co r d ing the m_nimum c urr c nt re q uired £ u r d rive m o toroper a ti o n a nd determining the t o rqu e used fr o m a mot o r cal ibration c urve ofminimu m c urrent versus torque. P ropell er thru s t for the tests was determinedby s ubtra c ting t h e drag mea s urement obtained whi le the propeller w a s operatingf ro m the drag of the c onfi g uration with the p r opel l er an d propeller s hroudremoved . The tu n nel-o p eratin g pr o peller te s t c ondition s are l isted intab l e I I .

Aco usti c T e sts

Me as urem ent s ys tem.- Fig ure 8 shows a s y stem blo c k d i agra m for a ty pi c a lmic r o p h on e c hann e l. T he p r i n c i p a l syste m c omp o n e n t s were a pre ss u re mi c r op h o newith a c ce s so r y n o se c on e , p r e am p li f i e r, p o we r s u pply, v a ri a bl e - g a in am pl i f i e r ,a nd FM t a pe r ec o r d er (o p e r a te d a t 76 . 2 c m/ s e c (3 0. 0 0 in. / s ec )) .

P ri o r t o th e win d- t un n e l te s t , th e syst e ms were a s se mbl ed a nd th e c ri t i c a lp ar a meters o f fr eq uen cy r esp on s e , di s t o rti o n , line a rit y , a nd e l ec tr o ni c noise

f lo ors were d oc umente d . These p rocedures and test results are summ a rized in

6



_}_ , i nserti o n in to t he preamplifier, ia sh ow n in f i gure 9 . The r o ll - off beginning_ at about 12 k Hz is c aused by the low-pass filter in the tape-recorder reproduce

e le ct roni c s. Daily pis to n-ph o ne l evel c al i b r a t ions were also made on site at a

level of 124-dB SPL at 2 50 Hz.

Be c ause t he analyses in this repo ct were based ma i nl y on A-welghted soundp ressure levels, no s y stem c orrec t ions have been included. The microphoneresponse was obtained f rom e lec t ros t atic labor a t ory c al i bra t ions, and t hesignal condi t ioning response was ob t a i ned fr o m the voltage i nser t ion sweepsdi s cussed previously. The nose-cone response c ame from m a nufa c turer's d a t a .The e f fec t of the wind t unnel on the recorded no i se da t a was obtained from da t ai n referen c e 4.

E leven mi c rophone sys t ems were used in t h i s tes t . The mi c rophones weremounted about the verti c al projection of the propell er hub on the tunnel a c ous-t i c ground board. The microphones were oriented so t hat the nos e cone was p a r-allel t o the free s t ream. A plan view of the airplane and microphone array is

shown in figure 6.

_ii_ ) ii_ From the photographs of the test configuration presented in f igure 4, itc an be no t ed that the airplane propeller axis was above the mi c rophone arrayp lane for e = 0o. Thus, t he acous ti c angle ( t he angle between t he hub-mi c rophone line and th e propeller axis) of each mi c rophone was no t t he same

as the angular location of the microphone with respe c t to the hub projection

_ on the microphone array plane. I ncluded in figure 6 a re both the array plan-view angle an d t he acous t ic angle for each microphone position wi t h the a i r-plane at _ = 0°.

:4 Data descri_ ion .- The data points listed in t able IV were chosen foracous t i c analys i s. These data were chosen t o prov i de t he w i des t possible vari-

L ation of power, thrust , blade angle, a nd noise at _ = 0° and zero sidesl i p.even t y se c onds of acous t ic data were recorded for mos t of these data poin t s.The tunnel-opera t ing ambient noise levels were de t ermined a t t est speeds ofabout 25 m / sa c (82 ft / se c ) and abou t 29 m / sac (95 ft / sec) with the airpl a nepropeller rem o ved.

Da t a acquired for t he conditions lis t ed in t able IV were analyzed to pr o -vide A-weigh t ed sound pressure levels and narrow-b a nd spe c tra (BW = 50 Hz).The A-we i gh t ed sound pressure levels were used t o pro v ide noise directivitypa t terns and t rend plots for each propeller conf i gurat i on, whereas the narrow-band spectra were used to show the detailed acoustic differences between pro-pellers and between sta ti c and tunnel-opera t ing conditions.

i_ RESUL TS AND DISCU SS ION

. : ... Aerodynamic Character istics

_I Effect of Reynolds number .- The results of the Reynolds n u mber tests with_ ii i _ pr o pellers'removed for a fl a p deflection of 00 are shown in figure 10. There

_ is a fairly l a rge effect of Reynol d s n u mber on the maximum lift and associated7

i

_;_I"



drag characteristics at test Reynolds number_ of ] 0 × ]06 and 2.1 x 106 but

lot: Reynolds numbers at and above 2.8 x ]06 , there is little effect of Reynoldsnumber. Most of the tests were therefore conducted for a Reynolds number of2.8 × 10 6 .

Effect of free propeller.- The effects of the 2-blade free propelleron the aerodynamic characteristics of the airplane are shown in figure 11.Increasing propeller thrust is seen to p_ovJde a small increase in llft-curveslope; for tests in which maximum lift was a chieved, the maximum lift wasincreased with increasing thrust as expected. In general, increasing thrust

resulted in a small increment of diving moment and resulted in slightlyincreased longitudinal stabili ty.

Effect of shrouded propellers.- The effects of shrouded propellers on air-pl a ne aerodynamic characteristics are given in figures 12 to ]4. These d a tashow the effects of thrust and blade angle for three different propellers. Thedata presented in figure 12 are for a 3-bl a de propeller with highly twisted(unloaded) tips. Figure ]3 shows data for the same 3-blade arrangement but

with a normal tip-twist distribution, and figure ]4 presents data for theunloaded-tip, 5-blade propeller arrangement. In e a ch case, propeller thrustand blade angle are seen to have very little effect on the longitudinal stabil-ity characteristics of the airplane.

Propeller Characteristics

The aerodynamic characteristics of the test propellers in forward flightare presented in figures 15 to 18, and the static-thrust characteristics of

the four test configurations are shown in figures ]9 to 22. All the propellerdata are seen to vary fairly uniformly with increasing blade angle, revolutionspeed and velocity (V / nD), and angle of attack. It should be noted t hat theefficiency data presented represent propulsive effioiencies (differences in

airplane drag with propellers removed or operating) rather than propellerefficiencies (shaft thrust). This actual shaft thrust may be masked somewhat

if the basic drag characteristics change with propeller operation since thrustherein is defined as drag measured on the scale system with propellers offminus drag measured with propellers operating.

Propeller Performance

Since one of the main purposes of the subject investigation was to deter-mine the relative merits of four different propulsive configurations, thethrust required was calculated for the airplane and plotted (along with theavailable thrust) as a function of flight speed in figure 23. The calculationswere made for full- and partial-power conditions for each of the four test con-

figurations. The thrust-required c a lculations were fo_ a wing loading W / S of957.6 N / m 2 (20 ft-lb) for s_andard sea-level conditions. The operating condi-tions assumed for the calculations were a s follows:

8

b_



lO0-pe reent powe r = 134 kW (180 hp) at 2700 rpm , variable I_

lO0-pe r cent powe r = 138 kW {185 hp) at 5000 rpm var iable B

C r uise s__ 9 _d__[qL j r e___ f ean fl _sl_ r ouded p ropeller -

Pb = 75 percent of full power and 80 percent of full revolution speed,variable 8

The results of the cal c ulations, along with maximum static-thrust valuesfor the four test propeller configurations, are shown in figure 23. It shouldbe noted here that all of the propeller c onfigurations have the same basic air-

frame and, therefore, the same "propulsive system off" drag. The shroud and

pr o peller of the shrouded arrangement are considered to be the propulsive

device that produces the net thrust for the propeller thrust coefficient,

It is readily apparent in figure 23 that the free propeller provides mu c h

more available thrust in flight than the other propulsors and that the 5-blade,

shrouded-propeller arrangement was the poorest. Maximum speeds of the free

propeller, 3-blade, normal-tip, 3-blade, unloaded-tip; and 5-blade , unloaded-

tip, shrouded propellers were, respecti v ely , 130 knots, 121 knots, 118 knots,

and 109 knots. It is also seen that unloading the shrouded-propeller tip wasdetrimental t o the thrust.

The airplane cruise performance varies in a l mo st the same manner as maxi-

mum speed conditions vary with maximum cruise speeds for the free propeller,the 3-blade, normal-tip , the 3-blade, unloaded-tip, and the 5-blade, unloaded-

tip, shrouded propellers of 115 knots, I01 knots, 98 knots, and 94 knots,

respectively. Ag a in, the free propeller was the best of the four arrangements

tested. The data also indicate large penalties in rate of climb for the

shrouded propellers.

An in t eresting p o in t can be noted for the static-thrust points shown in

figure 23( a ). The free propeller provided 3527 N (793 ib) of thrust, whereas

the 3-blade, normal-tip , shrouded propeller provided 3438 N (773 Ib) of thrus t .In other words, the smaller, 3 -blade, normal-tip, shrouded propeller did pro-

duce values of static t hrus t about equal to those of the free propeller , but at

forward-flight c onditions, the drag of th_ shroud arrange ment severely degraded

performance. It should be noted that the propeller-blade angles investigatedfor the 3-blade, normal-tip, shrouded propeller were not large enough to absorb

the available horsepower , so in o rder to obtain data c o mparable t o the other

configurations (i.e. , full rpm and Ps) , the data were extrapolated to largervalues of I_. The extrapolation showed that a blade angle of about 21 ° would

be required, with a resulting FM of about 0.74. Another point is that unload-

ing the pr o peller tip reduced the static thrust.



lower, by as much as 10 dB , for the shrouded configuration. This r e sult may

be ca u sed by shro u d shiel d ing. Comparison of figllres 2 4(b) and 25(b) show thatthere is no differ e nce in the in-plane noise under tunnel-oper e ting _onditi o ns.

A-weir]ted SPL variation with _._9_we__[r.- comparison o£ the A-weighted soundpress u re levels for various propeller c onfigurations plotted against power f o rmicroph o ne posltion 2 is shown in figure 26. Data for b o th st a tic and tunnel-

op e rating conditions ar e sh o wn. The free-propeller noise levels a re generallylower than other configurations regardless of power or flight condition. This

conclusion is typical of the results from the other microphone locations.

Narrow-band acoustic spectra. - N a rrow-band spectra for microphone posi-tion _ are plotted in figures 27 a nd 28 for se v eral propeller configur a tions.Ba sed on the experimental and theoretical results published in references 5and 6, increased noise from shrouded propellers occurring at high tunnel speedsmay be caused by inflow turbulence. Reference 5 indicates that a multiplicityof harmonically related tones is generated when a rotor encounters inflow tur-bulence. These tones, combined with tradition a l rotational noise harmonics,result in a discrete acoustic spectrum at higher sound pressure levels and athigher frequencies than for n onturbulent inflow. In reference 6, data are pre-sented which indicate that propeller noise may be expected to decre a se whengoing fr o m static run-up to forward flight. The reason given for this effectis that, for a static run-up, atmospheric turbulent eddies are stretched by thecontracting flow into the propeller disk. This causes a long-period turbulencedisturbance and many acoustic blade-passage harmonics to be generated. In for-ward flight, these same eddies are not stretched to nearly the same extent asfor static run-up; thus, the period of turbulence ingestion reduc e d, and lowernoise levels are obtained. Unpublished turbulence data fro m the Langley Full- -Scale Tunnel indicate that turbulence levels increase with tunnel velocity andapproach ]0 percent. Tunnel turbulence characteristics are not representativeof those oc c urring in free air. The spectr a for the free and shro u de d propel-lers are compared directly for stati c a nd tunnel-operating c onditions in fig-u res 27 and 28, respectively. As shown in the tables on each figure, the spec-tra were obtained at maximum and as nearly equal power settings as _ossible.For every comparison , the free-propeller noise levels were generally as low as,or lower than , those for the shrouded propeller for most of the power andthrust range investigated.

CONCLUDIN G REMARKS

Aerodynami c and a c ou st i c t e st s of f o ur differ e nt pusher-pr o peller c onfi g u-ra t i o ns on a t win-boom, general- a via t i o n airplane co nfigurati o n in t he LangleyF ull- Sc ale Tunnel sh o w t he following resul t s:

]. The free pr o peller had t he best propulsi v e c haracteri st i c s in both

st a ti c a nd f o rward-fl i ght c ondi t i o ns, followed ne x t by the 3-blade, normal- t ip,shrouded pro p eller, the 3-blade, unl o aded-tip pro p eller, and finally the5-blade, unloaded- t ip p ropeller.

11

• . . ...... __ , , . ,_ ,,,. ./_ __='- ,._,.T-_ -_-



<

I L 2. The free-propeller noise levels wore generally as low as, or lowerthan, those for the shrouded pr_pellers under tunnel flow conditions for most

i_i of the power and thrust range investigated•I 3 Statically, the propeller in-plane noise levels of the shro u ded

I propellers at high power conditions were less than for the free propeller.

At other than t he in-plane positions, the free- and shrouded-propeller noiselevelswere a bout equal•

Langley Research Center

National Aeronautics and Space Administration

Hamp t on, VA 236 6 5

Ja n u a r y 30, 1 980

REF ERENCE S

I. Me c h t ly, E. A.: The I nterna t ional System of Units - P hysical Constan t s

and Convezsion Fac t ors (Second Revision). NASA S P -7012, 197 3 .

2. Black, Donald M.; Wainauski, Hart: S.; Rohrbach, Carl: Shrouded

P ropellers - A Comprehensive Study. AIAA P aper No. 68-994,Oct . 19 68 .

3 . Hey s on, H a rr y H . : Use of S up er positio n in Di gi tal Comp u te r s To Obt a in

Wind-Tunnel Interference Factors for Arbitrary Config u rations, WithPar t icular Referen c e to V / S TOL Models. NASA TR R-302, 1969.

4. Abrahamson, A. L.; Kasper, P . K.; and P appa, R. S.: Acoustical Charac-

t er i s t i c s of t he NASA- Langley Full-Scale Wind Tunnel Tes t Se c t i on.NASA CR-132604, 1975.

5. Hanson, Do n ald B.: Study of Noise Sources in a Subsoni c Fan Using Measured

) Blade Pressures a nd Acoustic Theory• NASA CR-2574, ] 975.

6. Me t zger, Frederick B.; a nd Magliozz i , Bernard: New Directions in Air c raf t

P ropulsor Noise Research. [ P reprintl 750515, Soc. Automo t . Eng.,Apr. ]975.

12

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'I AI_,I,|.I . - _' I A'I I-'L' I. I_I'I'ARAI_I.I'I'I.,R:I

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TABLE I I .- TUNNEL-OPERATING PARAMETERS

JAngle-of-attack range, 0 ° to 16°; flap deflections of 0°, 20 ° , 3 0°I

Propeller Blade angle, rpm range Speed range,deg m / sec (ft, l sec)

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

Rem o ved --- 15 (49) t o 36 (]]8

2B, 1.981-m (6.50 ft) diameter 12 1000 to 2700 29 (95)16 800 to 2250

2 0 700 to 200 024 500 to 1800

.i_ 28 5 0 0 t o 15 0 0

3BT, 1.006-m (3.30 ft) d iameter 12 1400 to 5000 29 (95)_ , _,,_:. :., , ,: 16 1300 to 5000

__ _0 llOO t o 5000

24 1000 t o 450028 900 to 3500

32 800 to 3500

" ' 3B, l.O06-m (3.30 ft) diameter 12 1500 to 5000 29 (95)16 1300 to 5000

20 1100 to 5000

2 4 1000 to 450028 900 to 3500

32 800 to 350 036 700 to 3000

5BT, 1.006-m (3.30 ft) diameter ]2 1400 to 5000 29 (95)16 1300 to 5000

24 1100 to 4000

3 2 800 to 3500

36 700 to 3000

14



TA BLE III.- S UMMARY OF S YST EM LEVEL T EST S

Tes t P r o ced u re Tes t res u l t s, J , •........

Frequency response Apply os c illator signal at preamplifier +2 dB_ Thro u gh

inpu t . Record sys t em freq u ency resp o nse -] dB ] ]0 Hzthro u gh t ape-recorder o utp u t.

Distortion Apply signal a t micr o phone u sing aco u stic <2 percen tcalibrat o r. Check sys t em distortion

thro u gh tape-recorder o u tp u t.

Linearity Apply o scillator signal at preamplifier ±].0 percen t

input. Check system linearity at tape- of full-s c ale

recorder o u tput o ver expected range tape-recorder

se t t i ngs of variable-gain amplifier, deviation

Noise fl o or S h o r t c ir c u i t preamplifier inp u t and 4 0 t o 6] dB

(ref. 0.2 _N / m 2) monitor system noise level at tape-

recorder o u tp u t .

iI 15



. i

TABLE IV.- DATA POINTS CHOSEN FO R ACOUSTIC ANALYSIS

(a) St at ic c on dit io ns

................................................ C-I.........ul*nel R u nn Pt-o_x_ . 1 er Velocity , (_ , _[ , rp m Sideslip , Power , Thrust ,condition m / sec deg dog I d e g deg kW N

Static 20(05) 2B 0 0 10 0 1500 0 13 979

Static 20(07) 2B 0 0 10 0 2000 0 34 1802

Static 20(08) 2B 0 0 10 0 2250 0 5] 2335

Static 20(09) 2B 0 0 10 0 2500 0 75 2985

Static 20(]0) 2B 0 0 10 0 2700 0 98 3403

Static 21 (02) 2B 0 0 12 0 1 500 0 17 1085

Static 21 (03) 2B 0 0 12 0 2000 0 43 2033

Static 21 (04) 2B 0 0 12 0 2250 0 65 2656

St a tic 21 ( 0 5) 2B 0 0 12 0 2500 0 96 3360

Static 21 (06) 2B 0 0 12 0 2700 0 128 4014

St a tic 22( 0 2) 2B 0 0 14 0 150 0 0 21 12 63

S t a tic 22(0 3 ) 2B 0 0 14 0 2000 0 57 2 3 49

S tati c 22(04) 2B 0 0 14 0 2250 0 83 2 9 4 9

Static 22(05) 2 B 0 0 14 0 2500 0 127 38 3 9

Static 23(02) 2B 0 0 ]6 0 1500 0 28 ]379

Static 23(0 3 ) 2B 0 0 16 0 2 0 00 0 75 415

Static 23(04) 2B 0 0 16 0 2250 0 110 3132

Static 24(02) 2 B 0 0 18 0 ] 500 0 37 1450

Stati c 24(03) 2B 0 0 18 0 2000 0 98 2624

Static 25(03) 2B 0 0 20 0 1750 0 78 2055

Static 26(04) 2B 0 0 22 0 1740 0 96 2082

Static ]]6(02) 3B 0 0 12 0 3000 0 10 60g

Static 116(0 3 ) 3B 0 0 12 0 40 0 0 0 26 109g

btatic T 1 6(0 4 ) 3B 0 0 12 0 4500 0 3 7 1 3 92

Static 116(05) 3B 0 0 12 0 5000 0 53 1717

Stati c 125(03) 3 B 0 0 16 0 3 000 0 16 845

Static 125(05) ]B 0 0 16 0 4000 0 40 153q

) Static l _r-(06) . . , 3B 0 _ . 16 0 45 00 0 57 1 q 44

Stati c 125(07) 3B 0 0 16 0 i3000 0 7 q " .4 .I............................. _

%

16



TA BLE IV . - Con ti n u e d

(a) Co n c l u d e d

Tunn e l Run s Pr op e l le r V elo c ity , _ , 8 , _f , r p m S ide s lip , Po we r , Thru st ,co n d iti o n m/ s e c d e g deg d e g d e g k W N

S t at i c ] 26(03) 3B 0 3 000 0 24 ]1 2 1Stat i c 1 26(0 4 ) 3B 0 40 0 0 0 59 1 997

S t atic 1 2 6 (05) 3B 0 5 0 00 0 1 25 32 0 7S t at i c 95( 0 2) 3 BT 0 3 000 0 10 569S t at i c 9 5( 0 3) 3 BT 0 35 00 0 1 7 796

S tati c 95(0 4) 3 BT 0 40 0 0 0 27 106 8S tati c 9 5 (05) 3B T 0 4 500 0 40 1 3 1 7

S t a ti c 9 5 (0 6 ) 3 BT 0 , 5000 0 5 6 1624S tati c 9 6 ( 0 3 ) 3BT 0 30 0 0 0 1 6 8 0 5Stati c 96 (04) 3B T 0 35 00 0 2 5 10 8 ]

Stati c 9 6 (0 5 ) 3ST 0 4000 0 3 9 415Static 96(06) 3B T 0 4 5 00 0 5 7 18 2 8S t a ti c 96(0 7 ) 3B T 0 5 000 0 7 9 2260S t a t ic 104(02) 3 BT 0 3 0 00 0 2 2 1010

S ta ti c ]04(0 3 ) 3 BT 0 4 0 0 0 0 5 6 ] 79 3S tati c 104(04) 3B T 0 4 5 00 0 83 23 1 3

Stati c ] 0 4 ( 05) 3B T 0 5 0 00 0 ]16 291 8Static 105(04) 3 BT 0 0 _000 0 77 2273

Stat i c 105(05) 3 BT 0 0 4500 0 I] 2 292 7St a tic 64(05) 5BT 0 0 3000 0 11 592

S tati c 64(06) 5 BT 0 0 350 0 0 19 8 10Stati c 64(0 7 ) 5B T 0 0 40 00 0 29 1059S t a ti c 6 4 ( 0 8) 5B T 0 12 4 500 0 43 1 3 17

Stati c 64 ( 09) 5 BT 0 1 2 50 00 0 6 2 1 6 7 3

S tati c 70( 0 5 ) 5 BT 0 16 3000 0 21 83 6

Static 70(06) 5BT 0 16 3500 0 3 3 1090

S tatic 70(07) 5BT 0 16 4000 9 | 59 1503

17

, • ,. . ............. _---_ ..... I IIII I I i l



TABLE IV.- Continued

(b) Tunnel-operating c onditions

Tunnel Runs Propeller Velocity , (x , _ , i 6f , rpm Sideslip , Power , Thrust , V/ nDcondition m / sec deg deg i deg deg kW N

Flow 2(01 ) None 25 0 -- 0 .... 0 0 0 .....Fl o w 3 (01 ) None 29 0 -- 0 .... 0 0 0 .....

_> Flow 49(01) 2B 29 0 12 0 1500 0 42 258 0.565, _ Flow 50(0] 2B 29 0 ]2 0 2000 0 32 881 .433Flow 51(01 2B _9 0 12 0 2500 0 76 1837 .346

Flow 52(01 2B 29 0 12 0 2700 0 I05 2371 .324

Flow 53(0] 2B 27 0 12 0 2700 0 107 2486 .300Flow 54(0] 2B 25 0 12 0 2700 0 112 2678 .270Flow 46(0l 2B 29 0 16 0 1500 0 19 565 .570

Flow 47(01 2B 29 0 16 0 2000 0 47 1575 .442

Flow 48 (01 2B 29 0 ]6 0 2250 0 92 2126 .390Flow 30(01 2B 29 0 24 0 1500 0 48 1214 .575

Flow 31 (01) 2B 29 0 24 0 2000 0 I28 2393 .483Flow 37(01) 2B 29 0 20 0 1000 0 73 1392 .581

Flow 118(01) 3B 29 0 12 0 3000 1 0 6 147 .550Flow 119(01) 3B 29 0 12 0 4000 0 22 436 .4]5

<. Flow 120(0]) 3B 29 0 12 0 5000 0 48 867 .238

Flow 122(01) 3B 29 0 16 0 3000 0 12 294 .559Flow ]23(0]) 3B 29 0 16 0 4000 0 35 721 .415

Flow 12 4(01) 3B 29 0 16 0 5000 0 76 1357 .3 40

_;I) Flow 128(01) 3B 29 0 20 0 3000 0 II 463 .558 ;

Flow ]2gCo] 3B 29 0 20 0 4000 0 53 1050 .420Flow 130(01) I 3B 29 0 20 0 5000 0 114 ]877 .3 4 1

Flow ]32(01) 3B 29 0 24 0 3000 0 2g 649 .560

Fl o w 1 33 (0] ) 3 B 29 0 24 0 4000 0 78 1414 .425

Flow ]34(0]) 3B 29 0 24 0 4500 0 117 ]939 .37g

18

I,

ii.,

P ,"



TABLh IV. - Conclud ed

(b) Concluded

Tunnel Runs Pt'opel le t Velocity, n , _ , '_t', rpm Sideslip , Power, Thr u st / lid

cond i t ion m/ see de9 1 d o g deg deg kW N

Flow gl (0l 3BT 29 0 |12 0 3000 0 7 147 0.555

Flow 92(01 3BT 29 0 |12 0 4000 0 2 2 436 .422Fl o w 93(01 3BT 29 0 |12 0 4500 0 34 618 .375

ii_.;, F low 9 4(01 3_T 2_ 0 |1:' 0 5000 0 50 5 41 .335,_ .: :) F low 97 (01 3t_' 29 o 116 0 20 00 0 12 276 .'_5_' Flow 98(01 3BT 29 0 1 6 0 4 0 00 0 34 694 .420

Flow 99(01 3BT 29 0 116 0 5000 0 76 1272 .335

Fl o w 100(01 3BT 29 0 120 0 3000 0 18 423 .558

Flow 101 (011 3BT "_q 0 120 0 4000 0 51 q,"_- ._ •420

_ 4500low ]02(011 3BT .9 0 |20 0 0 75 1317 . 3 78

Flow I03(01) 3BT 2g 0 120 0 5000 [ 0 i11 1761 . 3 40

Flow I07(011 3BT 29 0 I24 0 .I000 0 28 600 .558

Flow 108(01) 3 BT 29 0 1 2 4 0 4000 0 75 13 03 .420Flow 109(01 ) 3BT 29 0 24 0 4500 [ 0 111 1757 . 3 74

Flow I l 1 ( 0 1) 3BT 29 0 28 0 3 000 ] 0 38 769 .560

Flow 11 2 (01) 3BT 29 0 28 0 40 0 0 0 100 1646 .42 q

Flow 66(011 5BT 2 9 0 12 0 30 10 0 8 165 .556

Flow 67(01) 5BT 29 0 12 0 40G" 0 25 4 6 3 .420

Flow 68(01) 5BT 29 0 1 2 0 4500 0 40 658 . _76

Flow 69(01) 5BT 29 0 i2 0 5000 0 57 85 q .340Flow 72(01) 5BT 29 0 16 0 3000 0 14 280 .552

Flow 73(01) 5BT 29 0 16 0 ._.500 0 24 485 .481Flow 74 (0l) 5BT 29 0 16 0 400 0 0 40 654 .421

Flow 75( 0 11 5BT 25 0 16 0 4 500 0 58 588 ..180

Fl o w 7 7 (01 ) 5BT 29 0 lb 0 5000 0 84 I.h18 . _45

Flow 82 (01) 5BT 25 0 2,1 0 3000 0 33 b32 . 160

Flow 8 . 1 (01) 5BT 2q 0 24 0 ,1000 0 8,1 141 q ..126

Flow 86(01) 5B T 29 0 3 2 0 1'i00 0 104 1704 .454

Flow 8q (01 ) 5BT 25 0 I¢_ 0 I000 0 52 ) I_8 _ .568.................................................... - .[.......................



22

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25



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26

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,, -- .... c:r .... __'_---'_2 2 712-_r -_- - " : .... _-- .......... .. . ........................... IN liJ n i



(b) 3-blade , twisted-tip propeller.

Figure 5.- Continued.

[ 2qIJ

t,

'z

, .... ,,,-.. --_ ..... , ....



.32

d/R

"2Lt 16

•20 12

t/d • 16 8

•08 0

0 -80 ,2 ,Lt .G .8 t ,0

r/R

(c) 3-blade, normal-tip propeller.


3O



t2

•32 \

•28 \ \

d l R .2u_

d l R _

\ -.20

/ o, °eg-- 13,deg

•12 k -8X

t/d

•08 t/d. _ - ,._ -12

•04 16Spinnersurface

0 -20O .2 ._ .6 .8 t .O

r/R

(d) 5-blade propeller.

Figure 5.- Co n cluded.

31



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36



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';_ 0 wnmlg---_--- 0.029

_ .283

1.2 _ 1_ l ,

_.o __ _ _ • I /8 ' , 'I' l I ,

% _ l _ Ii II/r

/ ' ,1 ' i ,

I

.2 J

.o _-__= _

.,A I.2 _ -"'"_L"-"_ _ .._ ..- - _ .... ,,.1 1

-._ I I-_ 0 _ 8 12 .2 0 -.2 -. 3 -.2 -.1 .0 .1 .2

u, a _ Cm CD

(a) _ = 1 2°

Figure 11.- Aer o dynamic chara ct eristi c s o f airpla n e wi t h _re e propeller operating.

_f = 0o; NRe = 2.84 x 106 .

37



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am ---i ....... -_i_ ..... 0 wnmlq_ "

"'_ ........ [ __--i..... _ ---_, --- o.o21............... --- _--- .2._0

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I i_ _ _ _ .07 0i . ------,4---- .1 02' _ L. i _ '"_- .... 214

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0 wnmlg

C_ i_ii_. ._l -- ---E)--- 0.004o -- _-- , 039.....(, , I 1 _ _ 092....... I_ _- 164

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¢ //_ ._ ,t" L L : i I/

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CO i' -- I Iu , .l

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/ ,,-_ 0 4 B 12 . 2 0 -.2 -.2 - .1 .0 .l . 2

"_ " C[, d e _ C m C D

,:". Figure 13.- Conti n u e d.

49

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•2 0 wnmlg---- ---_--- 0.013

(Ira 4_'_I_. ---e--- .o59o "_ , __._ ----&---- .131

• _ _ _ . 230

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om .......Wno,,m-,, -__--_---0.022

_111 .172

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CI, d_g C m G D

(d) 13 = 24 o.

Figu r e 1 3 . - Co ntinued.

51



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----El----- 0.010Cm ----_i_ _,q _,, ,,,-. 0 .065o "_'_ _- ",_,-- . 1 08



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