FUNDAMENTAL INVESTIGATIONS OF AIRFRAME NOISE
1
M.G. Macaraeg
2
NASA Langley Research Center
Hampton, VA, USA
1
This paper is declared a work of the U.S. Government
and is not subject to copyright protection in the United States.
Abstract
An extensive numerical and experimental study of
airframe noise mechanisms associated with a subsonic
high-lift system has been performed at NASA Langley
Research Center (LaRC). Investigations involving both
steady and unsteady computations and experiments on a
small-scale, part-span flap model are presented. Both
surface (steady and unsteady pressure measurements,
hot films, oil flows, pressure sensitive paint) and off-
surface (5 hole-probe, particle-imaged velocimetry,
laser velocimetry, laser light sheet measurements) were
taken in the LaRC Quiet Flow Facility (QFF) and
several hard-wall tunnels up to flight Reynolds number.
Successful microphone array measurements were also
taken providing both acoustic source maps on the
model, and quantitative spectra. Critical directivity
measurements were obtained in the QFF. NASA
Langley unstructured and structured Reynolds-
Averaged Navier-Stokes codes modeled the flap
geometries excellent comparisons with surface and off-
surface experimental data were obtained.
Subsequently, these meanflow calculations were
utilized in both linear stability and direct numerical
simulations of the flap-edge flow field to calculate
unsteady surface pressures and farfield acoustic spectra.
Accurate calculations were critical in obtaining not only
noise source characteristics, but shear layer correction
data as well. Techniques utilized in these investigations
as well as brief overviews of results will be given.
Introduction
The importance of reducing subsonic approach
airframe noise has now become apparent to the
international community.
1
Civil air traffic continues to
increase as does pressure from the public to control the
resulting increase in landing noise which is particularly
annoying to those living in close proximity to airports.
It is clear that noise reduction technology is critical to
the future development and operation of the world’s air
transportation system.
In response to the need for acceleration and
augmentation in key subsonic technologies, NASA
initiated the Advanced Subsonic Technology (AST)
Program in Fiscal Year (FY) 1992. The NASA Noise
Reduction Program began in FY 1994 under the AST
Program and incorporated several key high payoff areas
critical to the development of a new generation of
environmentally compatible aircraft.
1
The Noise
Reduction Program established a goal of a 10dB
community noise impact reduction relative to 1992
subsonic transport technology. The goal will be
achieved by combined noise reduction improvements in
the engine system, the aircraft and its operations. In
FY95 the Noise Reduction Program began an intense
effort in airframe noise source reduction and this
research will be the focus of this report. NASA also
has a newly instituted base research program in
airframe noise under the Advanced Concepts to Test
(ASCOT) and Futuristic Airframe Concepts and
Technologies (FACT) Programs in FY98. All these
efforts will be critical in achieving the environmental
goals set by NASA Administrator Daniel S. Goldin “to
reduce the perceived noise levels of future aircraft by a
factor of two from today’s subsonic aircraft and by a
factor of four within 20 years.”
2
NASA’s airframe noise effort under the AST
Noise Reduction Program involves partners in industry
and academia. NASA Langley Research Center’s
(LaRC) role is to determine fundamental noise source
mechanisms by relating sound generation mechanisms
to fundamental fluid mechanics. This is a critical need
of our industry partners, since past predictions are
largely based on empirical data that no longer suit the
newer classes of aircraft. LaRC couples its building
block experiments and computations to full
configuration and large-scale tests carried out at NASA
Ames Research Center (ARC) as well as flight data
made available by Boeing. In addition, Lockheed
Martin is a third partner performing computational
aeroacoustics for experimentally defined sources. The
on the flap-edge noise source which began in 1995. Our
objective is to obtain direct information regarding the
actual noise generation mechanisms responsible for the
spectrum produced by the flap-edge flowfield.
Technical Approach: Components
Since the mid-70’s researchers have found that the
primary airframe noise sources emanate from the high-
lift system and undercarriage of subsonic aircraft.
3,4
Depending on the type of aircraft, the dominant source
vacillates between flap, slat and gear. Since none of
these components are designed with aeroacoustics in
mind, it is no wonder that their very structure gives rise
to noise. In an effort to begin a thorough study of these
complicated sources, a systematic investigation was
initiated at NASA LaRC, to look at the details of the
meanflow surrounding these aircraft structures.
Components (i.e., flap and slat) were first tested in
isolation with simple models that gave rise to the same
dominant acoustic source maps that were seen in full
configuration tests and at larger scale. The models
investigated were 3-element, unswept, partial span flap
configurations. In both the flap and slat flowfields
large-scale coherent structures were seen to dominate
the flowfield. Details of these complicated vortical and
separated flow systems were investigated with
advanced experimental and computational tools; giving
rise to finer scale studies of the fluid mechanics and
acoustics which could potentially play a role in airframe
noise generation. The work to be discussed focuses on
the flap-edge noise source.
Following detailed investigations of the steady
flowfield both experimentally (on- and off-surface
data)
5
and computationally (Reynolds-Averaged
Navier-Stokes),
6,7
flowfield fluctuations were measured
using hot wire and hot films. The steady flowfield was
also investigated for sources of unsteadiness using
numerical simulations with the RANS base flow
8
as
well as linear stability which pinpointed dominant
frequency ranges of unstable flow disturbances.
9
These
efforts guided the correlation of acoustic measurements
with proposed noise sources. Surface unsteady pressure
measurements helped characterize the signature of the
source followed by microphone array technology which
obtained both quantitative spectra and farfield
directivity.
10,11
Maintaining a simplified model, tests were
performed at flight Reynolds number to determine the
robustness of both the fluid mechanics and acoustics as
this important parameter was increased. These tests,
performed in NASA LaRC’s Low Turbulence Pressure
Tunnel (LTPT)
12
also enabled for a range of velocity
sweeps at constant Reynolds number so that reliable
sources were tried so that cause and effect could be
both understood and substantiated.
Experimental Studies
Experimental investigations guided the more
detailed computational work covered in a subsequent
section. The most in-depth studies of both the fluid
mechanics and acoustics of the flap-edge flowfield was
conducted in the LaRC Quiet Flow Facility (QFF). The
airfoil was a NACA 63
2
-215 Mod B wing (16-inch
chord, 36 inch span) with a 30% chord half-span
Fowler flap. This geometry was also tested in the
NASA Ames 7x10 wind tunnel ,
13
a non-anechoic
facility, which utilized the model at about twice the
size.
Initial investigations of the flap-edge flowfield
involved laser light sheets and oil flows.
14
Laser light
sheet images revealed a dominant vortex in the vicinity
of the flap-edge. However, a short video constructed
from the laser light sheet data did not reveal significant
vibration of this structure (a noise generation
mechanism originally conjectured). The signature of the
vortex track on the flap edge surface was captured by
oil flow applications on the pressure and suction side of
the flap and main element. Fig. 1. shows these oil flow
patterns.
5,14
The curved streamlines seen in both the
suction and pressure surfaces give evidence of the flap-
edge vortex. In addition, a smaller focus of streamlines
on the flap side edge much closer to its trailing edge
indicates the presence of a second, smaller vortex.
Further substantiation of the double vortex
system are clearly seen in 5-hole probe studies
performed in the QFF.
7
The measurements shown in
Fig. 2 are normal planes of vorticity on the flap edge
taken from these studies. The dual vortex system is
clearly seen. The downstream planes indicate one
dominant vortex resulting from merging of both
vortices. Note that at the trailing edge, the vortex is far
removed from the flap surface.
Acoustic maps of high intensity noise on the
flap side-edge closely mirror the fluid mechanics of the
flow. Two microphone array systems were developed
at NASA LaRC to quantify these results.
11
A large
aperture array using 35 microphones was constructed to
obtain high resolution noise maps. This array possesses
a maximum diagonal aperture size of 34 inches. A
unique logarithmic spiral layout design was chosen for
the targeted frequency range of 2-30 kHz. In addition,
a small aperture array, constructed to obtain spectra and
directivity, complemented the larger design. This small
array possesses 33 microphones with a maximum
diagonal aperture size of 7.76 inches. It was easily
moved in both azimuth and elevation about the model
mounted in the QFF. Custom microphone shading
40 kHz with an overall targeted frequency range for the
array of 5-60 kHz. Both of these arrays were used with
the NACA 63
2
-215 Mod B wing model described
above. In Fig. 3., source localization maps from the
large aperture array chart the progression of the hot spot
for frequencies from 5 to 20kHz.
11
At the higher
frequencies, this hot spot is localized on the edge of the
flap. This is consistent with the primary vortex grazing
the edge of the flap. As the frequency decreases the hot
spot moves downstream and inboard, while the merged
vortex system comes over the edge impinging on the
upper surface. Recently reported noise reduction
schemes by NASA Ames
15
indicate that flap-edge
fences, which increase distance between the vortex
system and the surface, can achieve noise reduction.
A test conducted in the LaRC LTPT on a
second model of a part-span flap was co-investigated by
NASA’s High-Lift Program element and LaRC’s
Airframe Noise team. This model, known as the Energy
Efficient Transport (EET) model
16
has a vastly different
cove design as well as camber on a flap optimized for
high-lift. The LTPT test allowed Reynolds numbers to
range from 3.6 to 19 million based on chord. A further
advantage of this pressure tunnel was the range of
velocities achievable at constant Reynolds number, an
important aspect for obtaining accurate scaling laws.
This was the first test performed with an acoustic array
at these high Reynolds number conditions. Array
development
17
and data acquisition was performed by
the Boeing Commercial Airplane Company.
Although
the primary acoustic sources present in the QFF/7x10
experiments appeared in the EET study, an additional
hot spot at low frequency was also seen off the trailing
edge of the flap. The acoustic image map indicating this
additional source is given in Fig. 4b, in contrast with
the map of the previously observed source shown in
Fig. 4a. Details of the flowfield that corroborate this
finding are discussed in the next section.
In virtually all studies conducted an abrupt rise
in noise intensity occurred following an increase in flap
deflection.
10,18
It was found both experiementally
5
and
computationally
7
that both flap deflections also induced
vortex bursting on the flap-edge system. The correlation
of this event with noise is ongoing.
Computational Studies
An intensive investigation of the fluid
dynamics associated with the high-lift system began in
FY 1995. Mirroring the discussion above, these studies
aimed at elucidating the flowfield surrounding the part
span flap.
Initially, computations focused on the model
tested in the QFF and 7x10 facilities. Validation of a
Reynolds-Averaged Navier-Stokes solution is given in
6
7
19 20
obtained in the 7x10 experiment compared against
several of these computations. Most of the studies were
performed using CFL3D on a structured mesh utilizing
the Spalart-Allmaras turbulence model. An excellent
comparison of Pressure Sensitive Paint (PSP) and
computations of the vortex signature on the suction
surface of the flap are given in Fig. 6.
7
PSP was the key
measurement technique for obtaining details on the
edge of the flap, and was used to validate the
computational findings. However, oil flow methods
were utilized on the flap edge
14
surface as well. Figure
7 compares oil flows and computational streamlines for
a highly loaded flap setting. Note the accumulation of
oil at approximately 2/3 chord. The separation and
attachment lines abruptly end in this focal point. The
computations clearly show that downstream of this
point the flow reverses and moves upstream.
7
The
most dramatic proof of the vortex bursting which gave
rise to this focal point, occurred when Particle Image
Velocimetry (PIV) was performed in the LaRC Basic
Aerodynamic Research Tunnel.
21
Remarkable
agreement with the calculations was obtained. Fig. 8
depicts streamwise velocity plots from RANS and PIV.
Note the large region of reversed flow where the vortex
lifts off the surface while undergoing bursting.
The above studies led to a series of
systematic investigations utilizing the RANS as the
model meanflow for both linearized stability studies
9
and temporal simulations.
8,22
The frequency ranges,
calculated from the linear stability analysis, exhibited
the highest growth rates; the numerical simulations
indicated that these frequency ranges also had the most
explosive growth. The latter calculations were
interfaced with a Lighthill Acoustic Analogy, obtaining
the first noise calculations for this complicated
geometry.
22
The hot spots obtained in the QFF seemed
to be explainable in terms of both cylindrical shear
layer and vortical instabilities as discussed in references
8,9 and 22. Fig. 9 shows clearly the strong disturbance
vorticity present in the shear layer that rolls into the
flap-edge vortex. The results were calculated from a
temporal DNS utilizing a normal cut of the part-span
flap RANS solution at approximately 50% chord. The
broadband nature of this instability becomes apparent,
as this disturbance remains significant from 5 to 30
kHz. However, at the lower frequencies the vortex
instability appears to be stronger than that of the shear
layer. Recall from Fig. 3 that the 5kHz noise map
indicates the maximum in intensity occurring inboard
of the flap-edge at the location where the vortex moves
rapidly over the edge and onto the suction surface (Fig.
2). The higher frequency noise maps in this figure
indicate maxima more along the edge where the shear-
layer instabilities should be dominant; this is confirmed
reader is referred to references 8 and 22 for more detail.
As alluded to in the previous section another
source appeared during acoustic measurements on a
second part-span flap model, the EET. Recall Fig. 4,
which indicates a high noise region just off the trailing
edge of the flap at low frequency. When the RANS
solution was performed for the EET geometry, a
dramatic difference in vortex trajectory was discovered,
relative to that seen in the 63
2
-215 flowfield. Fig. 10a
indicates planes of vorticity along the flap-edge of this
model. If one contrasts these cuts with that of the 63
2
-
215 flowfield (Fig. 10b) the vortex is seen to stay in
close proximity to the edge surface of the flap down to
the trailing edge for the EET flowfield. In contrast, the
63
2
-215 model flowfield has the vortex abruptly
leaving the surface by around 60% chord. By the time
the flow reaches the trailing edge in that case, the
vortex is far-removed from any solid surfaces. It is
believed that fluctuations in the vortex, being much
nearer to the flap surface and trailing edge for the EET
model, potentially give rise to the additional hot spot of
Fig. 4. Again, the acoustic signatures closely mirror the
findings of the fluid mechanics.
Conclusions
Fundamental studies of airframe noise sources
for subsonic aircraft are being conducted at NASA
Langley Research Center. The work presented deals
with the noise generated by the flap-edge flowfield. A
series of detailed experimental and computational
building block studies were described corroborating
several key noise source mechanisms associated with
this important component. Coordination of acoustic
source maps with Reynolds- Averaged Navier-Stokes,
linear stability analysis, and numerical simulations gave
rise to plausible noise generation mechanisms
stemming from both a cylindrical shear layer and a
primary vortex structure. Newly designed array
technology allowed for high Reynolds number testing
in a hard wall facility, along with more detailed
quantitative acoustic measurements in an anechoic
chamber. Calculations of highly accurate mean flows as
well as unsteady flow characteristics gave excellent
agreement with both fluid mechanics and acoustic
experiments of several part-span flap models. This
work is part of a larger effort in the NASA AST Noise
Reduction Program, which includes large scale testing
at NASA Ames and Boeing Commercial Airplane
Company as well as computational aeroacoustics
performed at Lockheed Martin.
Acknowledgements
The author would like to thank William Willshire,
Program Manager of the Noise Reduction Program for
responsible for results given: Dr. C. Streett, Dr. T.
Brooks, Dr. W. Humphreys, Dr. M.Khorrami, Professor
Geoffrey Lilley, Dr. C. Gerhold, Dr. D. Lockard, Dr.
K.Meadows, Dr. R. Radeztsky, Dr. W. Hunter, Dr. B.
Singer, Dr. M.Takallu, M. Sanetrik Dr. W. Anderson,
J.Underbrink, R. Stoker, and G. Neubert.
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a) suction surface
b) pressure surface
Figure 1. Oil flow patterns on the flap-edge of the 63
2
-215 Mod B wing.
