Flap Edge Aeroacoustic Measurements and Predictions

30 Pages·2000·1.9 MB·English

by  Brooks, Thomas F.

AIA A
AIAA 2000-1975
Flap Edge Aeroacoustic
Measurements and Predictions
Thomas F. Brooks and William M. Humphreys, Jr.
NASA Langley Research Center
Hampton, VA 23681-0001
6th AIANCEAS
Aeroacoustics Conference
June 12-14, 2000 / Hahaina, Hawaii
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics
1801 Alexander Bell Drive, Suite 500, Reston, VA 20191

AIAA-2000-1975
Flap Edge Aeroacoustic Measurements and Predictions
Thomas F. Brooks*
William M. Humphreys, Jr.'t
NASA Langley Research Center
Hampton, Virginia 23681-0001
ABSTRACT
models. Areas of disagreement appear to reveal when
An aeroacoustic model test has been conducted to the assumed edge noise mechanism does not fully
define the noise production. For the different edge
investigate the mechanisms of sound generation on
conditions, extensive spectra and directivity are
high-lift wing configurations. This paper presents an
presented. Significantly, for each edge configuration,
analysis of flap side-edge noise, which is often the most
the spectra for different flow speeds, flap angles, and
dominant source. A model of a main element wing
surface roughness were successfully scaled by utilizing
section with ahalf-span flap was tested at low speeds of aerodynamic performance and boundary layer scaling
up to a Mach number of 0.17, corresponding to a wing methods developed herein.
chord Reynolds number of approximately 1.7 million.
SYMBOLS
Results are presented for fiat (or blunt), flanged, and
round flap-edge geometries, with and without a0 medium speed of sound
boundary-layer tripping, deployed at both moderate and c flap chordlength
high flap angles. The acoustic database is obtained CN normal force coefficient with respect to c
from a Small Aperture Directional Array (SADA) of
CI, static pressure coefficient
microphones, which was constructed to electronically CO_ coherent output power spectrum of unsteady
steer to different regions of the model and to obtain far-
surface pressure with respect to far-field noise
field noise spectra and directivity from these regions. d distance from one sensor to another
The basic flap-edge aerodynamics is established by D directivity factor, Eq. (13)
static surface pressure data, as well as by dS(y) elemental surface area at y
Computational Fluid Dynamics (CFD) calculations and f frequency
simplified edge flow analyses. Distributions of f _,3 one-third octave band center frequency
unsteady pressure sensors over the flap allow the noise
Af spectrum frequency bandwidth
source regions to be defined and quantified via cross-
G,, auto-spectrum of noise measured by SADA
spectral diagnostics using the SADA output. It is found
G_ auto-spectrum of unsteady surface pressure at
that shear layer instability and related pressure scatter is
sensor
the primary noise mechanism. For the flat edge flap,
Ga..,. cross-spectrum between outputs of SADA and
two noise prediction methods based on unsteady-
surface-pressure measurements are evaluated and surface pressure sensor
i pressure sensor location number
compared to measured noise. One is a new causality
j 42f
spectral approach developed here. The other is a new
application of an edge-noise scatter prediction method. k acoustic wave number - to/a 0
The good comparisons for both approaches suggest that (I correlation length scale inchordwise edge
much of the physics is captured by the prediction direction
(3 correlation length scale in spanwise direction
*Senior Research Scientist, Aeroacoustics Branch, Associate from edge
FellowAIAA. L length of chordwise section that a sensor
_Research Scientist, Advanced Measurement and Diagnostics represents
Branch.SeniorMemberAIAA.
L' lift per unit span
Copyright ©2000 by the American Institute of Aeronautics and M. convective Mach number, U,./ao
Astronautics.Inc. Nocopyrightisassertedinthe UnitedStatesunder
Title 17.U.S.Code. The U.S.Government hasaroyalty-free license M, Av,; average M, see Eq. (8).
to exercise all rights under the copyright claimed herein for M0 tunnel Mach number, U0/a o
governmentpurposes. All other rightsarereservedbythe copyright n normal vector to surface at y
owner.
American Institute of Aeronautics and Astronautics
Reynolds number based on c and U0 cross-spectral phase between SADA and
_0(2, S
acoustic pressure time history sensor outputs
P_ surface pressure time history SADA azimuth angle, see Fig. 24
Fourier transform of p,, CO radian frequency = 2trf
Fourier transform of p._
INTRODUCTION
qc dynamic pressure based on convective speed
Uc Airframe noise can be dominant during airport
qo tunnel dynamic pressure approach and landing when the engines are at low
r distance =Ii_ power and the high-lift systems and landing gears are
r' effective source-to-observer distance for deployed. This becomes particularly true as present-
quiescent field radiation, see Fig. 19 day propulsive systems become quieter I. As a result,
r vector distance = x- y there has been an increased emphasis placed on the
SADA small aperture directional array measurement and modeling of non-propulsive
t time
components such as flaps, slats, and undercarriage.
inboard velocity along suction side
Usu
As reviewed by Crighton 2, a number of studies of
inboard velocity along pressure side
Upr
velocity at radius ro from center of vortex, see airframe noise were conducted in the1970's and ear!_y
UI,, 1980's. An early evaluation was performed by Hardin.
Fig. 9(c)
Empirical airframe noise studies and prediction
convection velocity
U_, developments include those of Fink 4 and Fink and
Uh hydrodynamic convection speed from the
Schlinker 5. A series of airfoil self-noise experiments
pressure to suction side, see Fig. 13
were performed by Brooks and Hodgson 6and Brooks
hydrodynamic convection speed over the
and Marcolini 7"8'9for trailing edge noise and wing tip
suction side, see Fig. 13
noise. The results of these studies formed the basis of a
flow velocity at edge on pressure side, Eq. (4)
Up comprehensive self-noise prediction method l° for
Us flow velocity at edge on suction side, Eq. (4) isolated airfoils. As part of a wing and flap high-lift
Uo tunnel velocity system, the flap is much more loaded aerodynamically
X chordwise distance from the flap leading edge than it would be if isolated. Because of this, it has been
X noise observer location vector found capable of producing much more intense noise.
Xp effective observer location vector for radiation Block II in wing, flap, and landing gear interaction
in quiescent field, see Fig. (19)
studies found flaps to contribute significantly to the
Y spanwise distance from the flap edge
overall noise. Kendall 12 and Kendall and Ahtye 13,
Y surface noise source location vector
using an elliptical acoustic mirror, found strong
,.y height above surface sensor, see Fig. 8 localized flap edge noise. This was confirmed by Fink
Of flap angle with respect to the main element and Schlinker 5 in component interaction studies.
edge convective-flow skew angle, see Fig. 8
Mclnerny et al. 14, Ahtye et ai. fS, and Miller and
7 circulation density
Meecham 16 performed cross-correlation studies
"_2,s coherence function, see Eq. (9) between unsteady surface pressures and noise field
F vortex circulation measurements for the tip region of an isolated wing,
6 boundary layer thickness single slotted flap, and triple slotted flaps, respectively.
60 boundary layer thickness at airfoil zero angle The side edges of the multiple flaps were found to
of attack significantly exceed other airframe noise sources 16
¢ coherence decay factor for g3
The 1990's produced an increase in airframe noise
q coherence decay factor for gl
research activity 17, particularly due to the NASA
0 angle between n and x, see Fig. 19
O' angle between n and x', see Fig. 19 Advanced Subsonic Technology (AST) program.
0 observer azimuth angle defined for Eq. (21) Several tests are particularly notable. A 4.7% scale
DC-IO aircraft model was tested in the NASA Ames 40
P medium density
by 80 foot wind tunnel, as reported by Bent et al. 18,
T retarded time, Eq. (i 2) 19 2021 •
noise transmission time from sensor to SADA Hayes et al. and Guo et al. " . Inflow m_crophones,
a phased-microphone array, and a parabolic mirror
0 SADA elevation (flyover) angle, see Figs. 5
directional microphone system were used along with
and 24
unsteady surface pressure sensors on inboard and
outboard flaps. The flap edge noise was found to
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dominatoethernoisesourcesS. ignificanctorrelationsQFF were presented by Meadows et al. 34.
werefoundbetweenedgepressureasndthemeasuredMeasurements included flap-edge noise-source location
noiseel.Noisereductiocnonceptwsereevaluat2e2d.A
mapping by a large directional (phased) microphone
seriesoftestsofalargeunswepwting(2.5ft.chord) array system, flap-edge noise spectra and directivity by
andhalf-spaFnowlefrlapwereconducteindtheNASA a smaller array, and cross-spectra between unsteady
Ames7x10footwindtunnela,sreportebdyStormest surface pressure sensors about the flap edge. Details of
al.", Horneetal.24,andStormsetal.25.Thetests
the microphone array design and methodology used in
providedbasicaerodynamdicataand,althoughthe the testing was presented by Humphreys et al. 35.
tunnewl ashard-walleldim,itedacousticwsereobtained
Microphone array testing methodology was refined and
usinglargephasedarraysof microphones.A quantified using the QFF systems, as reported by
computationastludyby Khorramei tai.26provided Brooks et al.36. The present study builds upon this
substantiaaglreemewntiththedata.Thiswasusedto work.
examintewopossiblenoisesourcemodelsn,amelya,
In this study, the generation and radiation of flap
vortex-instabilmityodealndashealrayervortex-sheet
model. edge noise for the flat (or blunt), flanged, and round
flap edge configurations are examined. The basic flow
Thepresenptapecroncernaswingandflapmodel pattern about the edge is studied using Computational
testedin theQuietFlowFacility(QFF)atNASA Fluid Dynamic (CFD) calculations and measured static
LangleyT. hemodeilsaNACA63_,-21w5ingwitha pressure distributions. Simplified flow calculations are
30%chordhalf-spaFnowlerflap.Thisisthesameas then developed to provide key aerodynamic parameters
thatusedintheaforemention7exd10footwindtunnel
needed for noise prediction and scaling. Cross-spectral
testatNASAAmes, except here the model is about one amplitude and phase between unsteady surface pressure
half the size. As reported by MacaraegL7 this model in sensors over the flap edge surface are analyzed to
the QFF has provided the means to closely examine the reveal the character of the hydrodynamic pressure field
aerodynamic and acoustic physics for slats and flaps. due to turbulent flow and the near-field flap-edge noise
Measurements of the flow field in the QFF, by generation. Coherent Output Power (COP) spectra
Radezrsky et al.27 included hot-wire, hot film, 5-hole diagnostics using the measured pressures and the noise
probe surveys, laser light sheet, and flap surface oil provide a measure of the noise source distribution along
flows. These measurements revealed a dominant flap the flap edge. The noise source thus determined is
vortex structure resulting from the merger of two examined for consistency with the _reviously
upstream vortices - one strong vortex, formed from the mentioned shear layer instability mechanism _6"3°'33. For
pressure side to around the flap edge, and a weaker the flat edge flap, separate noise prediction methods are
vortex formed at the flap side edge on the suction side. developed and validated from (1) a causality approach
In the vicinity of the trailing edge, the vortex is far that connects the noise to the cross-spectra between the
removed from the flap surface. Computational efforts surface pressure and far-field noise through
by Khorrami et al. 2s and Takallu and Laflin 29 using fundamental aeroacoustic formulations and (2) an edge-
Reynolds Averaged Navier-Stokes solutions (RANS) noise scatter solution. Both methods utilize the surface
duplicated the key mean features of the edge flow. pressure measurements on the suction and pressure
Streett 3° developed a computation framework for the sides near the flap edge. Next the noise spectra and
simulation of the fluctuating flowfield associated with directivity are presented for three edge configurations
this complex flap-edge vortex system. Streett's for different surface roughness, flap angles, and flow
computations, utilizing a calculated mean flow field 28, speeds. The spectra are then examined for scalability
further crystallized the shear layer instability and for each configuration using flap mean lift and
vortex-instability disturbance models 26 for noise boundary layer thickness descriptions.
production. Linear stability analysis determined
TEST SETUP AND METHOD
dominant frequency ranges of unstable flow
ulsturoa.nees "_l. _t.Juo "_"_ , •In a similar time frame,
The test model apparatus is shown mounted in the
followed with a semi-analytical and semi-empirical
Quiet Flow Facility (QFF) in Fig. 1. The QFF is a quiet
prediction model of this shear layer instability
open-jet facility designed for anechoic acoustic testing.
mechanism. Predictions from this model compared
For the present airframe model testing, a 2 by 3 foot
well with flap edge noise data when certain scale
rectangular open-jet nozzle is employed. The model is
parameters were used.
a NACA 632-215 main-element airfoil (16 inch chord
The initial aeroacoustic measurements for an and 36 inch span) with an attached half-span Fowler
instrumented version of the above model tested in the flap (4.8 inch chord). The flap is attached by an
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adjustablseet
of "U" brackets to minimize bracket
interference with the ideal flap flow field. The model is
held in place by vertical side plates, which are
themselves rigidly mounted to side plate supports of the
nozzle. In the photo, the model is visible through the
Plexiglas windows located on the side plates. The main
FlAP
element airfoil and flap are instrumented with static
pressure ports and unsteady pressure sensors 34.
A view of the main element and flap in the vicinity
of the flap edge is sketched in Fig. 2. The flat edge flap
is shown accompanied by edge modifications. When
_,_ RoundEdge
attached, the flange edge produces a cavity depth of I/8
in. The flange thickness is 0.05 in. The round edge
attachment is a half-circle cross-section shape that
matches the airfoil contour. The effect of surface
FIGURE 2. Sketch of flap edge treatments.
roughness on the flap edge noise was examined by
applying grit. For the flat edges, #60 grit at a density of
about 70 particles per square inch was applied on the
edge and both suction and pressure side surfaces over a
2 in. span. For the round edge, #120 grit at about 800
MAIN
particles per square inch was applied, but was restricted
ELEMENT OVERLAP
to one half of the round edge surface area - towards the
flap's pressure side. The intent of the grit was to
produce thickened and well-developed turbulent
boundary layers in the vicinity of the side edge. For
this paper, the main element angle was set at 16° and
two flap angles, _=29 ° and 39°, were tested. The gap
and overlap settings for these angles are shown in Fig.
3. The positions of the flush-mounted unsteady
pressure sensors in the flap edge vicinity are shown in
GAP: 0.0227 0.0231 _ _29 o
Fig. 4. The chordwise distance from the leading edge OVERL0A.0P2:,20.0132 \\ -\
isx and the spanwise distance from the side edge is y. _x
FIGURE 3. Flap and gap geometry.
As will be discussed, the sensors of present interest are
those on the pressure and suction surface. These
sensors are Kulite model LQ-34-064-5A. They are
aligned spanwise at .06, .81, and i.81 in. (sections A, B,
and C, respectively) from the edge. The chordwise
position for each sensor is given in Table I.
The far-field acoustics of the model are measured
by Small Aperture Directional Array (SADA), which is
seen in Fig. 1 to be mounted on a pivotal boom
positioned by rotational stepping motors. The SADA is
always 5 ft. from the center of the main element trailing
edge. It consists of 33 B&K l/8-inch microphones
projecting from an acoustically treated metal frame.
FIGURE 1. Test apparatus with SADA mounted on pivotal The aperture of the array is small, with a maximum
boom in QFF.
diagonal aperture of 7.76 inches. The small size
reduces bias error by locating all the microphones in the
4
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arraywithinapproximatethlyesamesourcdeirectivity,
regardlesosf SADA'selevationorazimuthposition
aboutthemodel.InFig.5,theSADAmeasurement _): -39"
[]
positionasredrawninasideview(oppositseidetothat
-56* (_: 56°
ofFig.1)ofthetestsetupT. heSADAisshownlocated [] in
MEAN
inaplaneperpendicutlaorandcentereodnthespanof -- SHEAR
-73* //// // LAYER 73 °
themodelc,orresponditnogzeroazimuthaanl gle(_ = rl
0°). ThepositionofSADAinthephotoofFig.1
correspondtosanelevationangle¢ = -124 ° in the -9[]0* iI __ / RE.FERAACNTED 9I'10*
-- RAY PATH
drawing. In Fig. 5, the SADA is seen positioned at _ =
107°, on the pressure side of the model. The open jet rl I"°"EL /!/ _ 1S0AD7A"
-107*
shear layer boundaries (defined at 10% and 90 % of the
potential core velocity) are shown as measured along [] []
-124° SlOE PLATE 124"
the _ =0°plane. A mean shear line is shown, which is
I-1
part of a curved three-dimensional mean shear surface 141"
defined mathematically from the shear layer
measurements. This is used in SADA processing to
determine shear layer refraction corrections. The
drawing illustrates the refracted noise ray path from the
flap edge source region to the microphone.
FIGURE 5. Sketch of test setup. The noise ray path from the
flap edge to the SADA is illustrated.
SUCTION EDGE PRESSURE Data acquisition and post-processing
___/ SIDE SIDE \_ The array microphones and surface pressure
sensors employed acquisition hardware consisting of
27• 23• 18 •42 •47
transient data recorders controlled by a workstation. All
26. 22• 17 36 •41 ,46
35 microphone channels (including 2 reference
16 35
microhones) were recorded with a 14-bit dynamic
21 • 15 34 .40
33 range, simultaneously with 32 pressure sensor channels
25• 14 32 39 45 using a 12-bit range, at a sampling rate of 142.857 kHz.
13 Two million 2-byte samples were taken for each
20• 12 31 acquisition. The microphone signals were high pass
1,. 11 x 3o filtered at 300 Hz. All channels had anti-aliasing filters
,38 ,,4\L -
set at 50 kHz, which is substantially below the 71.43
// I L,' I I"
kHz Nyquist frequency.
C B A A B C
Microphone and pressure sensor calibration data
FIGURE 4. Drawing of unsteady surface pressure sensor were accounted for in the post-processing. For the
distribution.
SADA microphones, regular pistonphone and injection
calibrations of amplitude and phase were made.
Amplitude and phase calibrations for the pressure
sensors employed a miniature speaker-driver capable of
Sensor CoordinatesT . high frequency output. The measured outputs were
# x(inch) # x(inch) #--4x(inch) / # x(inch) /
referenced to the output of a 1/8 in. B&K Model 4133
10 012 20 0.95 30 1_| 0.12 / 39 3.02
11 054 21 2.62 31 / 0.95 | 40 [ 2.18 microphone. (The high frequency outputs of the
12 I 0.95 22 3.45 32 J 1.78 | 41 I 1.35
present Kulite sensors are unfortunately limited. In this
I 13 137 23 3,99 33 / 2.20 ! 42 t 0.93
14 1.78 24 0.12 34 | 2.62 1 43 4.80 I
I 15 I 2.62 / 25 1.78 35 / 3,03 44 4.68 report, surface pressure spectral data is limited
: 16 3.03 26 3.45 36 | 336 45 ' 3.02
generally to 13.5 kHz, where flat frequency response
17 3,36 27 3.99 ! 37 | 0,(30 i 46 1,35
1 18 I 3.99 / 29 0.00 38 / 012 , 47 10.93 /
and signal-to-noise are good.) Initial post processing of
YA= 0.06 inch ys= 0.81 inch Yc= 1.81 inch the test data begins with the computation of the cross-
spectral matrix for each data set. The computation of
TABLE 1. Pressure sensor coordinates. the individual matrix elements is performed using Fast
Fourier Transforms (FFT) of the original data
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ensembleA.ll dataaresegmenteidnto 1000non- maximum Mach number of 0.17 for this model
overlappingblockseachcontaining212samples,
configuration, which corresponds to a main element
yieldinga frequencyresolutionof 34.88Hz. A
chord Reynolds number of 1.7 x 106. In order to
Hamminwgindowisused.
maintain attached flow on the flap, the boundary layer
Aconventionbaelamforminagpproacehm, ploying transition was fixed by serrated tape applied to the
matricesof cross-spectrabetweenthe array lower surface of the main element at 30% chord and on
microphon3e5s'36is,usedtoelectronicall"ysteer"the the leading edge of the flap. Pressure coefficient plots
arraytochosennoisesourcleocationsT. heprocessingrevealed very similar performance to the somewhat
accounftosrmeanamplitudaendphascehangedsueto larger Reynolds number conditions of the similar
model 23 tested in the Ames closed wall 7 x 10 foot
refractesdoundtransmissiothnroughtheshealrayerto
theindividuaml icrophoneosfthearrays.A mean tunnel. In the QFF, the flap angle with respect to the
refracterdaypathisillustrateidnFig.5.Thecorrection main element was o_= 29° and 39°, whereas the main
termsarecalculate3d3usingSnell'slawinAmiet's element was set at 16° and 20° angle of attack to the
metho3d7,modifiedtoaccountfora curvedthree- tunnel centerline. (Note that 16°, for the main element,
dimensionamleansheasrurfacedefinedintheshear is approximately equivalent to an angle of attack of
layer. about 5° in the closed wall tunnel.) The flap flow field
A key feature of the array processing is that
was found to be dictated almost entirely by the flap
spatial resolution (or sensing area over noise source
angle, which is measured with respect to the main
regions) can be controlled independently of frequency
element, and not the main element angle.
and steering-direction over broad frequency ranges.
The microphone shading algorithms methodology used For the present QFF testing, pressure and lift
is adapted from Refs. 38 and 39 and evaluated with distributions for the flap are presented in Fig. 6. The
respect to the present test in Refs. 35 and 36. Note that main element angle was 16°. The gap and overlap
for each test case, the cross-spectral matrix has a settings, shown in Fig. 3, differ only slightly from those
corresponding background matrix subtracted from it to of Ref. 27. Static pressure coefficient distributions at
remove extraneous system noise (measured microphone three spanwise locations of the flap are shown in Fig. 6
and sensor noise for zero tunnel flow speed). The array for the tunnel Mach number M0= 0.17 for the two a
processing references levels to an equivalent single values. The spanwise cuts are shown for y/c = 0.027,
microphone measurement. Spectra data are determined 0.208, and 1.875. The ratio y/c is the distance from
for each narrowband frequency (34.88 Hz resolution the flap edge compared to the flap chordlength c. At
bandwidth) of interest. Other spectral bandwidths that y/c = 1.875, at the center of the flap section, the
are presented in this paper are formed from the expected two-dimensional lift distribution behavior
narrowband spectra. with high suction peaks is shown for both angles. As
y/c decreases (meaning the flap side is approached),
the high suction peak at the forward (leading edge)
FLAP EDGE FLOW FIELD
stations are reduced and the pressure differential
diminishes. Near the side edge, a low-pressure region
In this section, the basic flap edge flow is
exists at a downstream section of the chord, which is
examined with respect to parameters required to
due to a strong vortex being formed on the suction side.
evaluate the unsteady surface pressures and related
noise field. Also shown in Fig. 6 is the normal force (normal to
chordline) coefficient CN, with respect to c, versus
Basic aerodynamics y/c. An additional y/c location of 0.625 is
represented here. It is seen that the sectional lift is
Extensive aerodynamic measurements for the
diminished as the side edge is approached except for an
present model have been reported by Radezrsky et ai.27.
increase very near the edge due to the presence of the
The model was shown to function as a high-lift device,
strong vortex on the suction surface. At the inboard
with the main element and flap properly interacting
station y/c = !.875, CN = 1.213 and !.567 for tx =
aerodynamically. The elements are close enough that
29° and 39°, respectively. The ratios of Cs and a
the flow acceleration about the leading edge of the flap
values show almost a linear dependence of lift to flap
significantly reduces the required pressure recovery at
angle.
the main element trailing edge, but the elements are
separated sufficiently so that the viscous boundary The vortex found on the suction surface near the
layers do not merge. This increases the overall lift, flap edge was shown in Ref. 27 to be a result of the
especially on the main element, compared to lifts strong primary vortex and a weaker vortex merging.
obtainable separately. The QFF facility produces a
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ylc= 0.2083 sheared-flow velocity across the pressure surface edge
ylc= 0.0271 4 which wraps around the vortex and "feeds" it.
4 __ cx=29 °
2 :', ___ o_=39 °
Cp Cp
0 0
-2 6 x/c _ -2 6 x/c
4 y/c= 1.875 2
2
cr
Cp
J
J
0
-2 6 _c _ 6 _c
_39 o
FIGURE6. Pressurecoefficientdistributionsandnormalforce
coefficientdistributionfortwoflap angles.
The primary vortex is formed along the pressure side
(bottom) edge and grows in size in the streamwise
direction, and a weaker vortex is formed near the
suction surface edge. Steady RANS computations of
,.,d.,7'
Ref. 28 found agreement with the basic measured
features of the merger of the dual vortex system and the
general location of the resultant vortex. For both the
experiment and calculations, the vortex bursts above the
suction side surface for the 39° flap angle case. This
bursting occurs when the local flow angularity is too
FIGURE7. CFD results of flap-edge-flow velocity vectors in
high or the axial velocity component is too low. Figure planesparallel toand.035inchesabovethesurface.
7 shows portions of the RANS solutions for the two
flap angle QFF test cases of the present study. The
contours show lines of constant static pressure on the
surface. Intervals between the lines correspond to Of primary interest for this study are flow
intervals in Ct, of .346. The two component vectors parameters that provide guidance in determining noise
shown are the calculated velocities over a projected sources and provide pertinent input to prediction theory.
surface defined at 0.035 in. (approximately a boundary If the flap edge noise problem is indeed an edge
layer thickness) above the suction and pressure scattering problem, one would view the boundary layer
character and associated velocities as primary
surfaces. Only the edge velocity vectors from the
pressure side are seen because of the oblique view of parameters. One should be able to tie these to surface
Fig. 7. The flow about the side edge surface is omitted pressure data to validate the noise source - somewhat
for clarity. The vector pattern clearly shows the similar in approach to that done in Ref. 6 for trailing
presence of the resultant vortex and its strong influence edge noise. We direct our attention to the edge pressure
sensors on the suction and pressure sides. These would
on the flap edge flow field. The vortex is trailed
downstream of the model, but the vectors show the be thc only sensors in the strong edge flow field and, at
tormation of the vortex is essentially attached at the top the same time, be in the near field of such a scattering
(suction) edge surface. The attachment is seen to be phenomenon. They should therefore be representative
just aft of mid-chord for the 29° flap angle case, but of the source region. Note that the flap side-edge
surface, between the suction and pressure sides, has
slightly forward of mid-chord for the 39° flap angle.
The vortex strength is mostly defined by the strong generally lower velocity and its sensors (#1 through #9)
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CFD Values (Simple Model Values) for Mo= 0.17 component of velocity McCos_, exceed the tunnel
I c( = 29 ° (x= 39 ° value of M0=.17. On the aft (downstream of mid-
SNeunmgboefr _in_ Mc 13cdeg. __, Mc l_c(leg. chord) suction side edge, where the attached vortex
.010 21 (.222) 77 (90) .020 .22 (.237) 82 (90) flow comes off the surface past the edge, the flow
.010 .21 (.222) 90 (90) .012 .27 (237) 82 (90)
1123 ._001030 ,._322 ((,22.2)222) i_ 8753 ((9900)) ..002550 .22695 ((..223377)) 6880 ((9900)) velocities are even higher, reaching up to about twice
14 .050 .24 (.222) : 82 (90) .050 .33 (237) 76 (90) the free-stream value. Forward on the suction side
15 .030 .35 (.285) 54 (51) ,045 .32 (.331) 43 (46)
15 ,040 .33 (.285) 48 (51) .045 .30 (.331) 46 (46) edge, the velocities are lower than those aft and the
17 .050 .31 (.285) 52 (51) .060 .28 (.331) 60 (46)
18 .120 ,25 (,285) 58 (51) .150 .24 (.331) 65 (46) cross-flow diminishes greatly with flow skew angle fl,
30 .010 ,13(.215) 28 (34) .025 .15 (.253) 10 (24)
31 020 .19(.215) 38/_{ .020 .205(.253128(24) approaching 90°. An unexpected result, to the present
32 .020 ,21 (.215) 35 .020 .205 (.253) 26 (24)
! 3334 .002530 .2222(.2115)2153332(l3 4):l .o02530 .22052(.253) 2255( 224)4/ authors, for the CFD flow field is the lack of anticipated
.030 .22 (.215 31 .030 .21 (.253) 25 (24) changes in t_ and M,. values with changes in flap
,030 .20 (215) 32 (34) .030 .205 (,253) 27 (24)
(8=.0453) I5=,0314) angle. Expected increases in M,. did not occur with
increased flap angle, even in regions further away from
the surface. It should be mentioned that Ref. 28 noted
TABLE 2. Calculated edge flow boundary-layer thickness and
velocity values. that the solutions, while remarkably good overall in
defining basic flow features, found disagreements with
measured velocities on the order of 10 to 15%.
Concerns about the thickness of the shear layer were
z also expressed. It was suggested in Ref. 28 that
improvements may be needed with regard to grid
Surface # Surface
resolution and turbulence modeling. Because of the
\ Sensor-_U(z) I _kSensor
importance of these parameters to the present effort,
t x
alternate calculations are made and are presented in the
following section. The CFD solution, however, is
_U© x ]3_Uc utilized in providing a reference for primary flow-field
Conv_-_ features.
v_i_
Simplified edge flow calculations
FIGURE 8. Flow above a surface sensor and an idealized
shedinstabilitywave geometryatedgeof flap.
Simple aerodynamic modeling is used here to take
into account Reynolds number and flap angle effects in
are not considered here as representative of the source
the definition of boundary layer thickness and velocity
region (although they are in the near-field of such
values. This complements the description of the
scatter). In Table 2, for the sensors indicated, values
complex three-dimensional flow field given by the CFD
are given for the near boundary layer thickness 8, the solution.
corresponding Mach number M,, and flow angle fl,,
determined over planes parallel to the surface and at From thin airfoil theory 4°,the sectional lift per unit
height z=& above the surface. The choice of _ is span L' equals
partially subjective. It corresponds to the outer edge of
c
the shear flow nearest the surface. An illustration of the L'
PUoF = pU o] _dx= qoCCN (1)
velocity field U(z) above an edge sensor is shown in
o
Fig. 8. The top view shown defines the angle tic from where p is the medium density and F is the airfoil
the normal to the edge. The subscript designation c
circulation given an incoming stream velocity of U0.
(for convective) is used to indicate the flow above the
The circulation density _ of the vortex sheet defines
sensors is assumed to also represent any moving
the airfoil in the stream from the leading edge at x= 0
disturbance or flow structure that may cause noise-
to the trailing edge at c, where c is the chordlength (of
producing pressure scatter at the edge. The
the flap in the present case). The dynamic pressure is
hypothesized convecting wake sheet illustrated in Fig. 8
qo = pU2o/2 and CN is the sectional lift coefficient
will be subsequently discussed.
defined by Eq. (I). Figure 9(a) shows a sketch of an
For both flap angle cases considered, Table 2 inboard section of the flap where the flow is essentially
indicates that &,M c, and fl,. remains generally two-dimensional. The velocity jump across the airfoil
invariant along much of the pressure side edge. The sheet is _(x) = u_,,- ut,r, where us, is the velocity
flow speed M,., as well as the cross-flow (or spanwise) along the suction side and Upr is that along the pressure
side. The mean or average velocity jump over the chord
American Institute of Aeronautics and Astronautics