ATIC aeroacoustic model rotor test at DNW

ATIC AEROACOUSTIC MODEL ROTOR TEST AT DNW

Atsushi Murashige

Hideaki Nakamura

Noboru Kobiki

Kenjiro Inagaki

Akihiko Tsuchihashi ,

Eiichi Yamakawa

ATIC (Advanced Technology Institute of Commuter-Helicopter, Ltd.)

2 Kawasaki-cho, Kakamigahara City, Gifu Pref., 504-0971 Japan

Abstract

This paper presents the major topics evaluated by

ATIC model rotor test performed at DNW from Jan.23rd to Mar. 6th, 1998. In this test, the performance and the aeroacoustic data were obtained for 3 types of blades (baseline blade and 2 advanced blades) under the level flight conditions ( g =0~0.34) and the descent flight conditions (g =0.12~0.2). The parametric study of HHC (Higher Harmonic Control) effect on the BVI (Blade Vortex Interaction) noise was also investigated. On the BVI condition at g =0.16, the flow visualization of the tip vortex was pelformed using LLS (Laser Light Sheet) and PIV (Particle Image Velocimetry) techniques. The blade deformation data was also measured by PGM (Projected Grid Method) with HHC-on/off cases. Based on the obtained results in this model rotor test, the effects of variable rotor rpm, number of blades, blade geometry and HHC on the rotor pel'formance and noise are discussed in this paper.

Notation Symbols

c Blade chord length=110. 7mm r Blade local span wise station(m) R Rotor radius =2m

Coordinate system centered at rotor center x positive in downstream direction

y positive in starbom:d direction

z positive in upward di.l'ection

Cp Pressure coefficient defined as ;

CT

C"

v

C = Prot~/ - PJtolrc

'

Yo

p(Dr)' '2 Thrust coefficient Torque coefficient

Wind tunnel speed (m/sec)

a Rotor angle of attack(positive in nose up direction)

g

~ ~ 1-li-IC

Advance ratio

Rotor azimuth angle (deg.)

HHC phase angle (deg.) which is defined as the rotor azimuth angle where the blade pitch angle becomes its minimum as below.

AC02. 1

Rotor rotational speed(rad/sec) p ai1· density

~

o 1 s o 3 o• ¢ ¢ HHC Abbreviations/Subscripts

BVI Blade Vortex Interaction HHC Higher Harmonic Control LLS Laser Light Sheet PGM Projected Grid Method PIV Particle Image Velocimetry OSPL Over all Sound Pressure Level( dB) SPL Sound Pressure Level( dB)

Introduction

Although civil helicopter has been strongly expected to play a very significant role in Japan where the terrain is mountainous and the plain is highly populated, the noise problem especially around the residential area and the concern for operational safety prevented helicopters fi:om being prevalent in Japan.

ATIC was established in 1994 to research and develop the technologies in aeroacoustic area to reduce helicopter external noise and in flight safety area. The major aeroacoustic research items which ATIC has been conducting are as follows;

a)prediction technique of rotor noise and performance (Refs.1,2)

b)ailfoil and tip planform (Refs.3,4)

c)design trade-study for optimization between noise and performance

d)HHC and active flap for noise and vibration reduction (Refs.5,6, 7)

e)variable rotor rpm system t)BVI noise analysis (Refs.8,9)

g)flow visualization techniques (LDV and PIV) (Refs.10,11)

To evaluate these research items comprehensively and to construct a data base for analytical code validation, ATIC conducted the wind tunnel testing at DNW in the beginning of 1998.

This paper presents the major results of this DN\V testing as the first report.

Objectives

The main objectives of the ATIC model rotor test at DNW are as follows.

(l)Evaluate the effect of the following items on rotor performance, acoustic noise and vibration in the \~oind

tunnel testing.

· Blade Geometry : Airfoil and Tip Shape · Number of Blades

· Variable Rotor Rpm ·HHC

(2)Data collection for aeroacoustic and aeroelastic codes validation by the wind tunnel testing

Test Description at DNW (l)ATIC Model Rotor System

The general specification of ATIC model rotor system is shown in Table 1. See Ref.l2 for more details . The rotor system was set in the open jet configuration as shown in Fig. I.

T bl 1 P a e nmary F ea ure o fATICM d lR t o e otor ::,ystem

s

Rotor blade dynamically scaled

2m radius,l10.7mm chord tip:baseline, type-! and

type-2

linear twist of -8 deg hub fully articulated

4 bladed and 5 bladed Rotor drive system hyd.l'aulic 200HP output Primary blade pitch collective:-6 to +18deg.

control @75%r/R

cyclic:-15 to +15deg. HHC amplitude:0.8deg. @105Hz

frequency:up to 105Hz phase shift:variable Signal conditioning 256ch pre-amp. on the

rotating frame

Data acquisition 256ch simultaneous system measuring and on-line data

processing capability data band width up to 20kHz

Operating condition rotor rpm:840 to 1050rpm max. thrust:800kgf max. rotor shaft torque:

140kgfm

Fig.! Model rotor setup in DNW

J

l - . - - - -l('l~D lrm.

~2.>'-1 9l% 9° rr· 5por J\o\1on :>6';1, SSY. 61!0 n.sY, ) e:·1o ) JS%

I I i I ' I ' I •,

-+

~

==--;L.,_-_-,_,_-__

~1.:---:-r --~,;_,-..j_+-

cbJi

I

I '"""" ''""'''_'"' _Cro~; Section Tyrre TYI'Ell I l'Yi'Ell I l'YrEU I l'Y'l'<:l !TYPE:!rYPEI! I ' I ' I '

ti"c ccnt~r of rotat,on (o;ee F•g. S) TYr~rJ ITPE!l TYP~i1:

Baseline blade ~---:'~"\len1m - - - . . ; ~r Type-! blade ' - - - -2[)<).1"""---~ r:r.l s~"' sc~''"" %~. ~~~ l>1'l'. s~ l<;. I'"''""'T"n'~""" Cro"S""'i""T11" (>O<F•i5! ! --- ---+----+-- --.~-' ' ' ' ' TYPE IT rfl'F. 0 Type-2 blade

Fig.2 Planforms of the model blades and span wise distributions of pressure transducers

(2)Model Blades

Three types of rotor blades were evaluated; baseline(rectangular). ! (BERP like) and type-2(ATI). Each of them has a 4 meter diameter with a nominal blade chord of 110.7mm and a linear twist of-8deg.

The airfoil section of the baseline blades is NACA23012mod. Both of tYJle-land-2 blades have the same airfoil section; an A.KlOOD tapering to an AK080A at the tip section outboard ofr=lGOO(mm).

Each of the three types contains one pressure instrumented blade with approximately 100 pressure transducers and two heavily stl'ain gage instrumented blades with about 20 strain gages to measure the blade load/deformations. The planform of each blade and the location of sensors are shown in Fig.2 to Fig.4.

(_ .• _ •. -_::.',.' _~ .... ::.·.·.·.-.. ·.··_''-.. ~ ..• ',.·~<:>-.. · .. ·---~;. ····--·;.... "-(., ") . --- ~·- --~·--->.)··-~:-..., 1'YPE l TYPEll TYPE Ill TYPEIV TY?EV n~:-····---~ ·tJ---.. - .... --- --u--Q. ___________ ··---.- · - - · ____ -::.:.·:~ ... --TYPEVl

Fig.3 Chordwise distributions of pressure traducers at span wise locations (common to 3 types of blades)

Sp•n ~\1\ion !1%> U% ~01\ ~0% St'1> 6(]';1. 7C'll

i

I I I I I I I

"71-~-rr-~-

-'lj'-

~-1

the centcc ol r::>tJticn Symbols

₀

flapwise gages

0

chordwise gage3

6

torsion gages

Fig.4 Span wise distributions of strain gages (common to 3 types of blades) (3)Test Case

A lot of test cases were carried out to evaluate the effects of the variable rotor rpm, the number of blades, the blade tip shapes and HHC. The test conditions at each test case are shown in Fig.5(a),(b) . The contents of each test are as follows;

(a)Acoustic and Performance

The acoustic and the pelformance data were measured under the various flight conditions including hover,lG level flight, and a sweep in the middle flight speed range.

For the rotor noise measurement, 27 microphones

AC02. :3

in total were used. The 12 microphones were set at the wind tunnel nozzle to measure the noise propagation in forward direction. The other 13 microphones were installed on the inflow microphone traverse system (maximum traverse range was x:::::: ± 4m) to get the carpet plot 2.3m below the rotor. The other 2 microphones were installed on both sides of the rotor test rig.

For studying the effect of number of the blades, only the baseline blade was used.

All the three types of the blades were used for the study of the variable rotor rpm effect, the thrust effect and the blade geometry effect,.

During this test period, the blade surface pressure data was also measured, which was required for the validation of the CFD software developed by ATIC.

(b)HHC

For the basic understanding of the HHC effects on the rotor noise and the vibration, the parametric study of the HHC input parameters was pedormed using the baseline blade in 5-bladed configuration. In this test, the pelformance data, the acoustic data and the vibration data were measured under the BVI condition of the descent flight. (V=33.4m/s, a =4. 72 o

CT=0.0064, Rotor RPM=I003rpm) (c)Tip Vortex Flow Visualization

The tip vortex trajectories of baseline blade and type-2 blade were measured by LLS. The flow velocity fields around the tip vortex was also measured by PIV.

This test was conducted on the BVI condition of the descent flight. (V=33.4m/s, a =4. 72° , CT~0.0064,

Rotor RPM=1003rpm)

In the baseline blade case, the HHC effects on the tip vortex trajectories and the tip vortex velocity fields were also studied by use ofLLS and PIV, respectively.

(d)Blade Deformation

The deformation data of the baseline blade in HHC on/off cases were obtained by the PGM method. The test condition was the same as above mentioned two tests.( (b) and (c))

15 10 'Ol5

•

.3 0 0

i\

u -5 -10 0.0 CTswcep 0.1 0 0 [ 0 AO HHC,P!V,L ,S,PGM

~~0

0~

~RPM

_'WZ:P

o___.~

/ '

/ 0.2

"

0 0

~~

0.3

Fig.5(a) Test cases (JJ.-a plot)

0.010

to

0.007 IIIIC.PIV

<>

I}PM'" t~) LLS,P<;f(

<>

I

<>~

\

<> /

/

<>

<>

~~ r,

<>

<>

<><>

0.009 0.008 0.006 _y 0.005 a !-'weep v 0.01 ~ 0 ;:::: 0.009

z

UJ

0

u:

u. 0.008 UJ 0 0 1-~ 0.007 a: ---TY:K TYPE-2

-/

,/-'·-·''

-~

<

'Y

-

o 8ASEUNE

.

o TYPE-1

-0 0.1 0.2

"

0.3 0.4 ~

)5:

~

11 TYPE-2 l ASELINE Fig.5(b) Test cases ( 11 -CT plot)

Test Results (!)Rotor Performance

Fig.G shows the level flight performance of three types of blades at the nominal thrust level (CT=O.OOG4).

As seen from this figure, while the baseline blade is the best in performance on the low wind speed

condition, the type-1 blade is the best on the high wind speed condition.

The type-2 blade shows the good performance all over the speed range, having both benefits of the baseline blade in the low speed region and the type-1 blade in the high speed region.

Fig. 7 is the CT -CQ curve of each blade on the high \vind speed condition (p. =0.32). From this figure, it can be seen that the performance of the type-1 and the type-2 blade is better than that of the baseline blade in this order at the lower thrust level. However, as the rotor thrust level increases, the difference of the performance among these three blades decreases.

The reason of this tendency is not certain at 0.0006

~

0.0005

l-ID

o.o004

6

u::

u. 0.0003 w 0 0 w 0.0002 :::l

c

~ 0.0001 1-0

-,

1/

TtPE-1 ~

~.

--

'

r-· ---

_...

~

--o BASCUNE o TYPE-1

1-'1YPE-2 ~

"

.

.

.

2

--v~

TYPE

I "

'Jf~ I

.

' \

I

I

0.006 0.0003 0.0004 0.0005 0.0006 TORQUE COEFFlCIENT (CQ) 0.0007

Fig. 7 Rotor performance of 3 types of blades on the high speed condition test condition: fJ. =0.32, "=-4.45' ,

R 0 =210m/s,CT=O.OOG4-0.0088 present. But, further investigation will be requested based on the blade surface pressure information for the understanding of this phenomenon.

(2)Variable Rotor Rpm Effect on Rotor Noise Fig.8 shows the spectra of the out-of-plane rotor noise at 3 rotor rotational speeds (859rpm, 1003rpm, 1050rpm) on the middle speed level flight condition (p. =0.16, "=-1.01). The rotor thrust level is maintained to 444 kgf in all rotor rpm cases. The microphone position is just below the advancing side of the rotor (x/R=O, y/R=0.9, z/R=-1.15). As shown in this figure, the rotor rotational speed has very little effect on the out-of-plane rotor noise, especially in the frequency

120

<Dtoo

3

I

_{859tpm (R Q=1 Bnn/s)} I

_I

) _{-- 1cm,.-n CR=mn!.l} I ' i: I I.

,I

· · · · 1(ffipn (R=n/s) .J UJ > _UJ 00 .J UJ a: 60

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fJ)

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UJ a: ₄₀ 0.. 0

z

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20 · 0 I

_'if;',,":

':\I;!~:.

.,

'

~~

)

'

I

'"tfB

i - _i

;.,

' )00

-"'

'"

'"

\01)1) 1\011 ROTOR RPM (rpm) _I 0 0.05 0. I 0. I 5 0.2 0.25 0.3 0.35 0

ADVANCE RATIO (I')

Fig.G Level flight performance of 3 types of blades test condition: J.L ;:::Q.-....,0.34, a: is set to calculated trim,

R 0=210m/s, CT=0.0064

10 100 100l

FRfQ.BIJf (Hz) Fig.8 Effect of rotor rotational speed

on the out-of-plane rotor noise

test condition: fJ. =0.16," =-1.01' , CT=0.0064 R =180-220m/s (859-1050rpm)

range under 1000Hz.

As for the broadband noise in the frequency range over 1000 Hz, the strong dependency on the rotor rpm can be recognized.

Fig.9 shows the variable rotor rpm effect on the spectra of the in-plane rotor noise. The test condition is the same as in Fig.8. The microphone position is the right-front of the rotor (x/R=-3.57, y/R=l.28, z/R=O).

As shown in this figure, the strong dependency of the in -plane noise on the rotor rotational speed can be seen all over the noise frequency range.

From these results about the out-of-plane and the in-plane noise, the noise reduction by the variable rpm can be expected only for the in-plane rotor noise.

800 &00 1000 1100

1--'f-4-++l+l--+++l+i+h

ROTOR RPM {rom)

oL---~~~~~~~~_L~~lW

10 100 1000 1<XXXl

~(HI)

Fig.9 Effect of rotor rotational speed on the in-plane rotor noise

test condition: "=0.16, a =-1.01" ,CT=0.0064

R Q =180-220mls (N=859-1050rpm) microphone :x/R=-3.57, y/R=l.28, z/R=O

(3)Number of Blades Effect on Rotor Noise

Fig.lO shows one example of the effect of number of blades on the rotor noise under the constant thrust conditions.

The OSPL of the out-of-plane rotor noise is

presented for the 4-bladed and the 5-bladed rotor as the function of the rotor angle of attack. The microphone position is just below the advancing side of the rotor (x/R=O, y/R=0.9, z/R=-1.15).

From this figure, we can see that the OSPL of the 4-bladed rotor is 2 to 3 dB higher than that of the 5-bladed rotor at the wide range of angle of attack. At higher angle of attack than 6 deg., however, no

AC02. 5

difference can be observed between the 4-bladed and the 5-bladed rotor.

~

co

"'C '--'

120

115

...J

110

a.

en

0

₁₀₅

100

.

.

.

.

_'~

~

.

. .

_~··

/

f..,_

'-.

- -o- • 4-BLADES

(',

--...-5-BLADES

I

I

-2

0

2

4

6

8

10

ANGLE OF ATTACK (DEG)

Fig.10 Effect of number of blades

on the out-of-plane rotor noise

test condition: I' =0.16, a =-1-8" ,CT=0.0064 R Q =210mls

microphone : x/R=O, y/R=0.9, z/R=-1.15

Fig.ll shows the carpet plots of mid-frequency band noise of 355Hz to 5.6kHz range for the 4-bladed and the 5-bladed rotor. These data are obtained by the inflow traverse microphone on the maximum BVI condition (I' =0.16, a =4" , CT=0.0064).

From this figure, the peak noise level and the size of the strong noise region are affected by the number of blades. The peak SPL of the 5-bladed rotor is about 2dB lower than that of the 4-bladed rotor in both BVI regions of advancing and retreating sides.

lwiND

T

355-5. 6kHz

( 1) 4-BLADES

(2) 5-BLADES

SPL (dB)

iii

130 Fl125

i:Jj

120

;;~

115 :110 ::: 105

1100

@ 95 K~

I ::

I

so

Fig.ll Effect of number of blades on the BVI noise test condition: I' =0.16, a =-1-8" ,CT=0.0064

R 0=210mls

microphone :inflow traverse micl'ophone (z/R=-1.15)

From these results, it is inferred that the improvement of the out-of plane rotor noise level by 5-bladed rotor comes from the lower blade aerodynamic load caused by the lower blade loading. Because the lower blade load has the possibility to reduce BVI noise by relieving the degree of pressure fluctuation on the blade during BVI. Further data evaluation is required to understand the disappearance of this effect more than a =Gdeg ..

(4)Blade Geometry Effect on Rotor Noise

Fig.l2 shows the in-plane rotor noise spectra of 3 types of blades measured by the nozzle microphone (x/R=-3.57, y/R=1.28, z/R=-0.38) on the low speed level flight condition (I' =0.08, a =-0.66' , CT=O.OOG4).

120 IWE..I I£ C01oo

:s

- -nPE--1 .... T\'f£-2 ...J UJ I

_'

>

eo

UJ ...J UJ 0: 60 ::> Ul Ul UJ ~

~~~~

~

,.

'"

0: ₄₀ a. 0 z ::> ₂₀ 0 !I)

3100

_{~ ~} ~ ~

<;]'---

;:;ril

~ ~ U) . . ' "

-<

,.

,.

'"

"'

""

0 10 100 1000 ~(Hz)

Fig.l2 Effect of blade geometry on the in-plane rotor noise on the low speed condition test condition: I' =0.08, a =-0.66' ,CT=O.OOG4,

R 0=210m/s

microphone : x/R=-3.57, y/R=1.28, z/R=0.38

1000)

As seen from Fig.l2, the noise level of the type-2 blade is the lowest all over the frequency range. The worst is the baseline blade and the type-! blade is in between.

Fig.13 shows the rotor in-plane noise spectra of three types of blades measured by the nozzle microphone (x/R=-3.57, y/R=1.28, z/R=-0.38) on the high speed level flight condition (I' =0.34, a =-5.05' ,

CT=O.OOG4).

As shown in this figure, the difference among each blade is very small. Because we expected the lower noise level would be observed in the type- I and the

type-2 blade case than in the baseline blade, this result is contrary to our expectation.

To understand the reason of this result, further investigation will be continued.

Fig.l4 shows the carpet plots of the mid frequency band noise of three types of blades, which is obtained by

the inflow traverse microphones on the descent flight

~110 <I t3100 !

I I

I

Ill

0

~

00

_{~==~~1IITIIIli[l}

=IJJIITill

10 100 1000 10000 FFm.B\C( (Hz)

Fig.l3 Effect of blade geometry on the in-plane rotor noise on the high speed condition test condition: p. =0.34, a =-5.05' ,CT=O.OOG4,

R 0=210m/s microphone : x/R=-3.57, y/R=1.28, z/R=-0.38

355-5. 6kHz

(1)

BASELINE

355-5. 6kHz

(2) TYPE-1

+WIND

355-5. 6kHz

(3) TYPE-2

:·: 110 -::J 105

gj100

:~

1

95

I

so ; 85 !II 80

Fig.l4 Effect of blade geometry on the BVI noise test condition: p. =0.16, a =4. 72' ,CT=O.OOG4

condition (IL =O.lG, a =4.72' ,CT=O.OOG4).

The upper plot is the carpet plot of the baseline blade, the lower left one is that of the type-! blade and the lower right one is that of the type-2 blade. AJ;, seen from this figure, the position and the size of the high SPL region is influenced by the blade geometry.

In the type-! and the type-2 blade case, the advancing side high SPL region is observed at the smaller azimuth position than in the baseline blade case.

And the size of this high SPL region is decreased on both blade cases compared to the baseline blade case. To understand the reason why the blade geometry

'

0

_,

-·

_,

_,

-10 I

\

I-:1YPE2 0 0 0

,,

_{• """ · : BASELINE} ~ f--.

'

oo

'

-50 -40 -31) -2.0 -10 0 10 20 30 40 50 ZIG(%)

Fig.-15 Tangential velocity distl'ibution in the tip vortex of 2 types of blades test condition: IL =0.16, a =4. 72' ,CT=0.0064 R 0=210mls

no HHC

HHC phasc=30dcg

p:::

-

Q) '-'

c:

"'

...

"'

·~

0

"'

"'

·~

::;:

0.2 0.1 0 -:TYPE 2 0 - o - : BASELINE ~

"

-0.1 -0.2 0 0.2 0.4 0.6 0.8

Fig.lG Blade/vortex miss distance of 2 types of blades

1.2

r/R

test condition: IL =O.IG, a =4.72" ,CT=O.OOG4 R 0=210m/s

affects the BVI noise, we try the PIV and the LLS to the baseline and the type-2 blade.

Fig.l5 shows the velocity distribution of the tip vortex measured by PIV and Fig.16 shows the blade/vortex miss distance measured by LLS.

From Fig.l5, we can see the tip vortex strength of the type-2 blade is stronger than that of the baseline blade contrary to our expectation.

From Fig.l6, the miss distance of type-2 blade in the blade tip l'egion, where the large BVI noise is generated, is lager than that of the baseline blade.

The type-2 blade has the lower noise in spite of higher vortex strength than the baseline blade. This can be explained that the type-2 blade has a larger miss distance at the blade tip region than the baseline blade has. We think the future work for evaluating the test data with various test condition will give us the sufficient insight for quantifying the effect of blade geometry on the rotor noise.

40dcg

50dcg

Frequency band : Mid frcq=355 to 5,600Hz

Fig.l7 HHC effect on BVI noise in the flow traverse microphone SPL carpet plat Glrev HHC, amp.=0.4deg

(5) HHC

The effect of HHC on BVI noise reduction is evaluated applying 6/rev HHC to the 5 bladed rotor.

The carpet plot of the mid frequency band (355 to 5600Hz) containing BVI noise obtained by the inflow traverse microphone is shown in Fig.17.

The high SPL regions caused by BVI noise propagation can be observed on both advancing and retreating sides in each 1/J H!IC case as well as no HHC

case. But the BVI noise is reduced by HHC with

cf; t!llt'= around Odeg.

In order to evaluate HHC effect in more detail by the sound pressure spectrum, 1lic#3 at the position

illustrated in Fig.l7 is selected as reference.

Fig.l8 shows HHC influence on mid frequency SPL represented by over all value of !Vlic#3. There can be seen a sinusoidal characteristic of OSPL with respect to HHC phasing. Some ~ HHc's alleviate BVI and others deteriorate. This figure shows that 1/J HHC

=Odeg is most effective on BVI noise reduction, which reduces the noise by about 3dB from that of no HHC case, and cjJ IU-!C =30deg is the worst.

iii

_::?,117

0'

116

"115

!

114

~

113

'\; 112

..J

111

Bi

110

0 n.

1/

0

"

-

-

'no HHC ,..( '0, -<r--HHC on

/

\

!'-..

rl

_{.. ""'-n}

' '

20

40

60

HHC phase (deg)

Fig.18 HHC effect on BVI noise in mid frequency OSPL of!Vlic#3

The sound pressure spectrum comparison among the best HHC case ( ~ HHC =Odeg), the worst HHC case

(~ IIIIC =30deg) and no HHC case are shown in Fig.19.

It is noted that the two HHC-on cases, even the best HHC case, have larger SPL than no HHC case in less than 1000Hz. This is caused by larger loading noise generated by HHC operation. But this SPL magnitude

120 110 100 iii 90

"

~ ~ ₈₀ ~ 10 50 50 100 1000 10000 Frequency (Hz)

order changes over 1000Hz range so that the

magnitude order of SPL becomes consistent with the OSPL as shown in Fig.17 and Fig.18.

Fig.20 shows the unsteady blade surface pressure characteristics with respect to the rotor azimuth comparing with and without HHC cases. The pressure data presented here come from 4%c position of the upper sulface at G blade span stations fi·om 67%R to

95%R.

The no HHC case has the severe p1·essure

fluctuation around t/J =60deg and 300deg, which causes

larger BVI noise as shown in Fig.17. But on the best

HHC case for BVI noise reduction ( ¢ !II!C =Odeg), the

pressure spikes on the advancing side are almost vanished. This ch:mge ln the pressure characteristics between the two cases makes BVI noise reduction.

noHHC Min noise HHC(6/rcv,ph~st.=Odcg)

-1.6 -t .4 -1.2 -1.0 c. -0.8 C) -0.6 -0.4 -0.2 0.0 0.2 0 180 270 360

Rotor azimuth Rotor azimuth

Fig.20 HHC effect on blade surface pressure

at 4%c on upper surface

#5

r

Vortex#J.

,#4

#}

#2

jfJ

/ r

~~·!Jr

... .

#4

i.

II ,

;1

#5

.

,.,,

..

';··'

/'

Wind

.

0_/l ~ ::i.

---o--

no HHC HHC on 6/rev phase=Odeg 0 " " (mm) Blade at <!J =60deg X

'""'

~r---+---, 0 •••. -o

""

Vortex#2

"·

_"·

""

·a.

_,.,

-=

r

-mL---'

Fig.21 HHC effect on tip vortex trajectory with the

The blade tip vortex trajectory obtained by LLS is shown in Fig.21 with the blade deformation measured by PGT\11 comparing no HHC case and the minimum noise HHC case(¢ '""' =Odeg).

Comparison between these two cases indicates that HHC application has a small influence on the horizontal trajectory of the tip vortex, but makes a large difference in the vertical trajectory projected on r-z plane at ¢ =GOdeg. Observing the trajectory of Vortex-#2, which is just before the interaction with the reference blade at cjJ =60deg, it can be seen that no HHC case has the interaction at about 80%R and the minimum noise HHC case has at about 72%R.

It is inferred that this inward movement of the span wise interacting position makes the difference in the degree of the smoothness of the Cp vs. c/J curve

shown in Fig.20 because of the difference of the spanwise blade local load between no HHC case and HHC on case at the interacting position. Then, the relief in SPL on the BVI region is made as shown in Fig.17. This inference will be verified later by evaluating the blade load distribution obtained by the pressure integration.

Fig.22 shows HHC influence on the rotor performance on BVI flight condition represented by Cq and HHC phasing of 5/rev HHC as well as 6/revHHC. The increment in Cq is more than 5% on 5/rev HHC with 0.7deg. amplitude and more than 1% on 6/rev HHC with 0.4deg. amplitude.

In both cases, the extra rotor power consumed by HHC is not fatal comparing to the rotor power of no HHC case. Furthermore, the flight condition which causes BVI is descending, where the required rotor power has a wide margin up to the engine available power. Consequently, it is predicted that HHC operation will

not cause a significant rotor power penalty.

--o--6/rev 0.4deg 0.0002000 0.0001950 0.0001900 0.0001850 C! _0.0001800 0 0.0001750 0.0001700 amp. HHC

•

5/rev 0.7deg

,...

amp. HHC

-

-

'no HHC

J

• 1

•

_·~

n 0.

l.n..

.~

...

-

-

-

-;;

•

0.0001650 0.0001600 0 20 40 60 80 HHC Phase (deg)

Fig.22 Influence of HHC on the rotor performance

Conclusion

(!)Sufficient data to evaluate ATIC's noise reduction techniques had been obtained in this wind tunnel testing.

(2)A part of data study for the effects of blade geometry, variable rotor rpm, rotor configuration and HHC on pelformance/noise is pelformed.

(3)Further effort for test data evaluation should be kept going.

Acknowledgments

The authors wish to express their thanks to Dr. Mercker and many DNW/NLR staff for their support to perform this \\lind tunnel test campaign and to the PIV team from DLR lead by Dr. Ehrenfried. Also our thanks to Mr. Archinal and Mr. Ingalsbe of DEI for their assistance to handle the model rotor system.

References

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Tsujiuchi, T., Yamakawa, E., Aoyama, T., Saito, S., " Comparison between Calculated Rotor Noise and Experimental Data obtained by DNW Test", 24th European Rotorcraft Forum, Marseilles, France, September 1998, Paper AC06.

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