Content uploaded by Lakshmi Sankar
Author content
All content in this area was uploaded by Lakshmi Sankar on Aug 12, 2015
Content may be subject to copyright.
GTRI Report: A5928/1
Application of Circulation Control Technology
to Airframe Noise Reduction
Annual Performance Report
NASA Grant NAG-1-2146
GTRI Project A5928
Prepared by:
K. K. Ahuja, L. N. Sankar, R. J. Englar, Scott Munro, and Yi. Liu
GEORGIA INSTITUTE OF TECHNOLOGY
Atlanta, Georgia 30332-0844
Prepared for:
NASA Langley Research Center
Hampton, VA 23681-2199
15 February 2000
ii
FOREWORD
The Georgia Tech Research Institute (GTRI), a unit of the Georgia Institute of
Technology, contracting through the Georgia Tech Research Corporation, is pleased to submit
this performance report to NASA Langley Research Center on “Application of Circulation
Control Technology to Airframe Noise Reduction” conducted under Year one of NASA Grant
NAG-1-2146. Any questions concerning this report should be addressed to Dr. Krish K. Ahuja
(Tel: 770-528-7054, email: krishan.ahuja@gtri.gatech.edu).
iii
TABLE OF CONTENTS
Section Title Page
1.0 STUDY OBJECTIVES AND THE TECHNICAL APPROACH..................................1
1.1 Study Goal ........................................................................................................1
1.2 What is Circulation Control?...............................................................................2
1.3 Use of Circulation Control to Reduce Airframe Noise......................................2
1.4 Method of Approach...........................................................................................2
1.5 Report Outline......................................................................................................3
APPENDIX A- CFD STUDIES OF THE CCW FLOWS……………………………A-1
APPENDIX B - A DRAFT COPY OF A DOCTORAL THESIS PROPOSAL ON THE
TOPIC OF THIS REPORT……………………………………………….…………..B-1
1
SECTION 1
STUDY OBJECTIVE AND THE TECHNICAL APPROACH
1.1 Study Objective
The objective of this proposed program is to explore the feasibility of utilizing pneumatic
or Circulation Control technology for reducing airframe noise.
1.2 What is Circulation Control?
Circulation Control (CC) is a recently-maturing pneumatic aerodynamic technology that
offers significant improvements over
the well-known jet flap and blown-flap
concepts of blown lift augmentation. In
the CC concept, the jet sheets or the
tangentially-blown flaps of these two
concepts are replaced by the non-
moving round or near-round CC
trailing edge slot or slots, Figure 1.
The blown sheet remains attached to the CC curved trailing edge by a balance between the
negative pressure differential across the jet and the centrifugal force acting on the curving jet.
The resulting flow entrained into the curving jet sheet initially acts as a boundary-layer control
(BLC) at very low momentum (blowing) coefficients to prevent separation. At slightly higher
blowing, the jet adheres to the round trailing edge, moving the airfoil's stagnation point and
streamline well onto the lower surface and acting as a pneumatic circulation control. This
greatly augments the airfoil lift well beyond that of mechanical conventional high-lift flap
systems, and into the region of Supercirculation (i.e., well beyond BLC). Because this
dramatically changes the airfoil static pressure distribution, the concepts can also be used to
modify the aerodynamic moments as desired, as well as increase/decrease the drag and the
downstream wake.
Figure 1a Circulation Control concept
2
1.3 Use of Circulation Control to Reduce Airframe Noise
Although the noise benefits of Circulation Control have never been shown by anyone
through experimentation, it has high potential for reduction of airframe noise and is the subject
of this study. This study was motivated by the expectation that pneumatic circulation control
wing (CCW) like devices for high lift operation will reduce noise by significantly reducing the
impact of flow separation, high angle of attack operation, large complex components exposed in
the freestream such as jet/flap interactions, trailing edge and flap edge vortices, large wakes
behind the airfoil, etc., associated with conventional high lift devices of today. Also, short
ground roll distances plus steep approach and climbout offered by this technology can
dramatically reduce the ground noise footprint. In addition, this method offers considerable
weight reduction. This will reduce the fuel requirements per flight, thus reducing the emission
impact.
One could question: What about the noise produced by the blowing jets used for CCW?
Fortunately, the width of the jet slot used for Circulation Control even in a full-scale system is
expected to be of the order of a fraction of an inch. This will produce dominant noise in the high
and ultrasonic frequency region and will have little noise impact on community noise. Also,
short distances plus steep approach and climbout offered by this technology can dramatically
reduce the ground noise footprint. In addition, this method offers considerable weight reduction.
This will reduce the fuel requirements per flight, thus reducing the emission impact.
1.4 Method of Approach
This investigation is to be accomplished through a highly integrated effort consisting of (a)
detailed aerodynamic measurements of lift and drag in a wind tunnel equipped with a 6-
component balance, (b) acoustic measurements in an anechoic flight simulation facility, (c) CFD
analysis of selected configurations.
A three-year program was originally proposed. This report is a performance report for
portion of the first year. A test configuration that has produced significant lift improvement and
has displayed a drag-control capability in our previous work was selected for noise testing in
Georgia Tech’s Anechoic Flight Simulation Facility. For the sake of comparison, another test
3
configuration consisting of an airfoil and a conventional flap has also been tested. A task on
CFD of pneumatic flow control was also initiated in Year 1. Based upon the results of the CFD
analysis and the initial acoustic tests, an optimized pneumatic configuration is planned to be
tested both for its aerodynamic performance and noise reduction capabilities.
The second year’s effort will concentrate on a similar investigation of pulsed blowing to
control airframe noise. The third year will be spent in investigating tangential blowing to control
trailing edge and flap edge noise.
Two doctoral students, Mr. Scott Munro and Mr. Liu, are working on this program. Mr.
Munro’s dissertation will concentrate on the experimental portion of the effort and he will be
advised by Dr. Krish Ahuja. Mr. Liu’s dissertation will concentrate on the computational effort
and he will be advised by Dr. L. Sankar. Both students will have Mr. Robert E. Englar available
for consultation on various aspects of the circulation control technology because of Mr. Englar’s
experience in this area for over 25 years.
1.5 Report Outline
Significant accomplishments have been made during the first year. However, to render the
reading of this report least painful, we have provided only summarized highlights in the form of
PowerPoint images in the main text of this report, which appear in the next section. A detailed
progress of the CFD task appears in Appendix A.
Likewise, should the reader desire to obtain an extensive background of the circulation
control technology and a description of the experimental results in more detail, we have attached
the draft proposal for Ph.D. thesis of Mr. Scott Munro as Appendix B. It includes a description
of the previous work on circulation control technology as well as an extensive description of the
initial results that form the basis of Mr. Munro’s doctoral dissertation. These results are also part
of the majority of the experimental results obtained during the first year of this grant.
4
SECTION 2
KEY ACCOMPLISHMENTS TO DATE
Significant accomplishments have been made during the first year. This section includes a
summarized set highlights in the form of PowerPoint images.
Figure 2 provides an executive summary of the program with bullets for the historical
perspective of why circulation control technology type configuration needs to be investigated for
reduced airframe noise along with the planned approach, accomplishments and future plans. It
also shows schematics of the type of wing used in the present study.
Figure 3 provides a list of disadvantages of multi-flap conventional wing and the advantages
of a pneumatic circulation control wing (CCW). This figure also shows a photograph of an
actual full-scale Navy aircraft A-6 Intruder where the circulation control technology was
demonstrated to be a viable concept.
Figure 4 further elaborates the points made in Figure 3 and includes initial acoustic data
obtained from the present investigation. Clearly the spectra shown here point out that for the
same lift, reduced noise levels can be obtained for a system equipped with circulation control.
Figure 5 shows the initial results of our computational study. The wing model used for
experiments was modeled. Velocity vectors with and without pneumatic blowing are shown in
this figure with a superimposed forward velocity. It can be seen that without blowing, a well-
defined vortex shedding is obtained downstream of the small flap used in the circulation control
slot. This vortex shedding is absent with the blowing. This is reflected in our acoustic
measurements as seen in the typical noise spectra shown on the right hand side of the same
figure. Without blowing, a well-defined discrete tone associated with the vortex shedding is seen
at a frequency of about 1600 Hz. This tone completely disappears on using the blowing through
the circulation slot. This is a powerful result in that it indicates that not only can we potentially
reduce broadband noise associated with airframe spectrum, but even the vortex shedding noise
produced by various bluff bodies in a steam can potentially be controlled using circulation
control technology.
5
Additional details of the progress on the computational task appear in Appendix A.
Finally, Figure 6 provides photographic views of the reference conventional wing and the
circulation control used in the present investigation. The conventional wing was operated with a
gap in its flap which is representative of cut outs provided in actual wings due to structural
constraints or to prevent engine exhaust from impinging on an extended flap in the case of wing-
mounted engines. Typical noise spectral comparison for the two wings mounted in GTRI’s
flight simulation facility is provided on the right hand side of this figure. Spectrum of the
background noise obtained by operating the empty tunnel background is also shown. Data for a
flight velocity of 220 ft/s at a microphone located in the flyover plane directly below the trailing
edge at a distance of 11.5 ft from the slot center is shown here. The CCW flap angle was set to
30. The slot velocity was adjusted to provide the same lift as produced by the conventional
wing. The lift data was obtained in our earlier studies using the same two wings in another wind
tunnel equipped with a six-component balance. Clearly, the CCW shows significant noise
benefit in the main audio range of frequencies. Actually, even if the two spectra were identical,
CCW configuration should be considered better because of its simple design and light weight.
Considerable amount of acoustic data has already been acquired. After initial examination of
this data, it became clear that identifying individual noise sources in the entire CCW system as
tested will be difficult without examining the noise field of an isolated thin slot of a span
comparable to that used in the CCW itself. Much value will be added if individual noise
components can be studied and used to determine how much each part affects the noise of the
entire system. Possible individual contributors are the jet flow from the slot, the changed
directivity due to curvature, or perhaps the jet impinging on freestream flow. Although much
work has been done in the area of jets, no one has extensively looked at the properties of
extremely high aspect-ratio jets, similar to the CCW slot. In the current study, aspect ratios
range from 1,200 to 10,000. There has been acoustic work on rectangular jets, but the aspect
ratios have rarely been greater than 10. Thus, to study the unique characteristics of a CCW slot-
like aspect-ratio jets, a high aspect ratio nozzle (HARN) has been fabricated. The HARN will be
mounted in the Anechoic Chamber at GTRI where it will be used to study the effect of aspect
ratio on the jet noise. Acoustic measurements will be taken at several jet Mach numbers, slot
6
heights, and aspect ratios. Acoustic data will again be taken with several microphones located in
the acoustic farfield at several angles relative to the jet flow.
Attachments to the HARN will also allow for investigation of jet turning on the directivity of
the noise. A variety of curved surfaces will be fabricated to simulate the CCW cylindrical
surface that creates the jet turning. Acoustic measurements will be taken in a similar fashion to
the aspect ratio tests discussed above. This nozzle is shown schematically in Figure 7 and is
currently being tested for its noise characteristics in GTRI’s Static anechoic jet flow facility.
The HARN was fabricated at GTRI in order to perform jet noise studies on very high aspect
ratio jet flows. The primary focus is to provide a similar jet flow model to the CCW wing, in
scale and flow, separate from the other aerodynamic influences incurred in the CCW system.
Because of the interest in maintaining similarity between the HARN and the CCW being tested,
the HARN was designed to be 30" wide (see Figure 7). The HARN mounts to a round-to-
rectangular transition duct section (2.75” square exit) that mounts directly to the 4" opening on
the plenum of an available jet rig. It gradually contracts to the desired slot height in one
dimension and expands to the 30” width in the other. Both of these dimensional changes occur
over about 30 inches of axial length. The HARN was designed with exit plates that are
adjustable. This allows the slot to have potentially an infinite number of slot heights within the
bounds of fully closed to 0.25". For the current study, most of the heights will be similar to
those in the CCW tests, ranging from 0.003" to 0.020". However, if it is determined after
examination of the initial data that acquiring data from more slot heights will further our
understanding, more heights will be added to the test matrix.
The HARN has been fabricated out of aluminum with an intended maximum pressure of
about 25 psig, enough to produce a slot exit Mach number slightly over 1.2. The external surface
of the HARN is sloped towards the exit, so that the entrained flow is more nozzle-like rather than
a wall-jet. A curved attachment piece to provide a Coanada surface is currently being fabricated,
again maintaining the shape and size of the CCW system being tested in the Flight Simulation
facility of GTRI. It will be attached to the HARN via a piece that will allow it to change angles
similar to deflecting the flap on the CCW.
7
Results from both of these studies will be compared to ascertain whether or not the noise
from these unique aspects of the CCW technology can be predicted by existing jet mixing noise
theory or by some correction to existing theory.
Note that noise spectra obtained with HARN configuration for most operating conditions
is expected to independent of any potential internal noise that might limit understanding of the
data from the circulation control wing for some conditions. This is because the HARN will be
mounted on a jet noise facility that has been calibrated for its acoustic cleanliness and past data
from which has been used extensively by the aeroacoustics community to validate jet noise
theories. This data will thus also help us identify the bounds of the acoustic cleanliness of the
CCW test configuration and provide further confidence in that the CCW data is not contaminated
by any upstream noise associated with slot air supply tubes
8
l
History
– Airframe noise & engine noise
compete during aircraft approach
phase.
– High-lift devices contribute
significantly to airframe noise.
– No alternatives to mechanical, high-lift
devices investigated.
l Alternative
– Circulation Control Technology.
– Significant potential in reducing
airframe noise.
l Studies, Modeling and Assessment
– Conduct study using CCT to control
noise.
– Model flow physics.
– Conduct assessment of lift using
conventional wing and circulation
control.
l Accomplishments
– CFD analysis carried out with blowing
and without.
– Lift and blowing increased
concurrently.
– Noise Reduction obtained in audio
range
Figure 2: Application of Circulation Control Technology
to Airframe Noise Reduction
l
Future Plans
– Understand conditions which produce noise reduction for
same lift.
– Understand drag implications.
– Extend investigation to pulsed blowing and tangential
blowing.
9
• Light
• Simple
• Variable CD
• Less Noise
• Maximum CL = 4-7
• Reduced Ground Rolls & Noise Footprint
Pneumatic CCW
Figure 3: Comparison of Existing High-Lift Systems:
Mechanical vs. Pneumatic
Mechanical Multi-Slotted Flaps
• Heavy
• Complex
• Draggy
• Maximum CL = 3-5
10
Figure 4: Noise Reduction through Circulation
Control Technology
LI
F
T
C
O
E
F
FI
l Mechanical high-lift configurations can employ 3 or more trailing-edge flap
elements plus one or more leading-edge elements. They also employ many
tracks, actuators, and support mechanisms to deploy/retract the devices, plus
mechanical spoilers and ailerons. These devices are Noisy and cause
compromises to be made in terms of weight, structure, complexity, reliability,
wing sizing, and overall performance
l Clearly, a system that can do away with most of the mechanical
components, reduce the wake width, reduce noise significantly, improve
performance and yet keep the aircraft flight-worthy is needed. One way to
accomplish this is to replace mechanical components responsible for lift
by pneumatic or Circulation Control technology.
l Potential for noise reductions
is very high.
l shorter runway
l no nose droop
l Simple design, smaller wings,
less weight
l Less emission
l Recent GTRI work
shows that blown high-
lift Circulation Control
Wing (CCW)
configurations can
generate more lift due to
blowing at zero degrees
incidence than the
maximum lift
coefficient
demonstrated by
mechanical airfoils, but 30
40
50
60
70
80
90
100
0 2000 4000 6000 8000 1 104
Conventional wing noise vs CCW,
α
= 0
o
, V
~ 220 ft/s
SPL (Pref = 20 X 10 -6 Pa)
Frequency, Hz
Tunnel Background noise at 220 ft/s
Highly Improved Lift of a CCW Configuration
l Circulation
control is an
innovative flow
control
technology that
can dramatically
improve
aerodynamic
performance and
simplify
mechanical
complexity
through
pneumatic
means.
l
Circulation control technology has previously been developed and
flight-demonstrated for civilian and military aircraft (for example,
A-6/CCW).
Conventional Wing
CCW, 30 deg. flap
11
Figure 5: Suppression of Vortex Shedding Noise
by Blowing
30
40
50
60
70
80
1000 2000 3000 4000 5000 6000
Trailing edge flap noise, V = 240 ft/s, CCW 30 degree flap, α = 0 o, h ~ 0.003
90 deg. microphone, ~11.5 ft. (∆ f = 2 Hz)
(8-26-99, junk, junk1)
SPL (P ref = 20 X 10 -6 Pa)
Frequency, kHz
CCW, 30 deg. flap
No Blowing,
Cµ
= 0.0
Blowing, Cµ
= 0.025
Flow Field around the Airfoil Cµ=0 Case
Flow Field, Cµ
= 0.15
12
Figure 6: Noise Reductions at Constant Lift
(Conventional Wing vs. CCW Wing)
Conventional Wing
30
40
50
60
70
80
90
100
0 2000 4000 6000 8000 1 104
Conventional wing noise vs CCW, α = 0 o, Voo ~ 220 ft/s
90 deg. microphone, ~11.5 ft. (∆ f = 32 Hz)
(9-4-99 g4, 8-25-99 T4, 7-15)
SPL (Pref = 20 X 10 -6 Pa)
Frequency, Hz
Tunnel Background noise at 220 ft/s
CCW, 30 deg. flap
13
30"
30"
Anechoic
Chamber
Plenum
Round to
Rectangular
Section
Hich Aspect Ratio Nozzle
(Top View)
2.75" h ~ 0.000" - 0.25"
Sliding knife edges
to make adjustable slot
Side View
30"
Flange fabricated to be
compatiable with existing
4" round to 2.75" square
nozzle section
Figure 7a: Schematic of high aspect ratio nozzle (HARN).
Figure 7b: Side view of HARN with CCW flap.
Appendix A
A - 1
APPENDIX A
CFD STUDIES OF THE CCW FLOWS
A unsteady three-dimensional compressible Navier-Stokes solver was developed. This
solver is capable of handling isolated wing-alone configurations. Both finite wings and 2-D
airfoils may be simulated with the same solver. The boundary conditions have been coded in a
general form so that the researcher may specify the slot location, slot size, blowing velocity, and
the direction of blowing. The effects of turbulence are modeled using either a Baldwin- Lomax
eddy viscosity model, or using the Spalart-Allmaras one-equation model.
A 3-D grid generation code was also developed that can model wings of general plan form.
Figure1 shows the body-fitted grid at a typical control station.
Figure 1. Body-Fitted Grid at a Typical Span Station for a Circulation Control Wing
Appendix A
A - 2
The flow solver was validated by computing viscous subsonic flow over a small aspect-
ratio wing made of NACA 0012 airfoil sections at an angle of attack of 8 degrees. The
freestream Mach number was 0.12, while the Reynolds number based on wing chord was 1.5 x
106. Surface pressure data for this wing are available from experimental studies done by Bragg
et all. Figure2 shows typical surface pressure distributions at three span stations. A good
agreement with measurements is observed.
Following the code validation, the 3-D Navier-Stokes solver was applied to the CCW
wing configuration shown in Figure1. Simulations were done both for a no-blowing case, and a
case were blowing was applied near the trailing edge, on the upper surface. Figure3 shows the
velocity vectors indicative of the flow pattern in the trailing edge region. For the no-blowing
case, a region of separation was present upstream of the flap, and downstream. The flow
downstream of the flap exhibited a periodic vortex shedding. With blowing, this recirculation
region and the unsteady vortex shedding were completely eliminated.
The primary benefit of the circulation control through blowing was in the lift coefficient.
While the wing in the no-blowing case produced a lift coefficient of 1.22, the same wing, with a
blowing coefficient Cm of 0.15, produced a lift coefficient of 2.6. Such as large lift coefficients
can presently be realized on conventional wings only with the use of high lift systems such as
flaps and slats.
In summary, it appears that the circulation control has two primary benefits: very large
lift coefficients without the use of high lift devices, (b) elimination of separation in the aft
regions of the wing. Both these benefits may be expected to have attendant noise reduction
benefits.
Appendix A
A - 3
34% SPAN
-3.5
-3
-2.5
-2
-1.5
-1
-0.5
0
0.5
1
1.5 00.2 0.4 0.6 0.8 11.2
CHORD
Cp
Upper- Exp
Lower -Exp
Lower Present
Upper Present
50% SPAN
-3
-2.5
-2
-1.5
-1
-0.5
0
0.5
1
1.5 00.2 0.4 0.6 0.8 11.2
CHORD
Cp
Upper Exp
Lower Exp
Lower Cal
Upper Cal
66% SPAN
-3
-2.5
-2
-1.5
-1
-0.5
0
0.5
1
1.5 00.2 0.4 0.6 0.8 1
CHORD
Cp
Upper Exp
Lower Exp
Lower Cal
Upper Cal
Figure 2. Computed and Measured Surface Pressure Distributions over an Unswept Wing
Appendix A
A - 4
Figure3 a. Velocity Field in the Vicinity of the CCW Wing Trailing Edge (No-Blowing Case)
Figure3 b. Velocity Field in the Vicinity of the CCW Wing Trailing Edge (Blowing Case)
Appendix A
A - 5
Work to be Done During the Next Reporting Period: The coordinates of an airfoil currently
being tested by Ahuja and Englar is shown in Figure4. The aft region of this airfoil may be
rotated, simulating an integral flap as in Figure1. During the next reporting period, a series of
aerodynamic calculations will be done for various flap settings, with and without blowing. The
blowing slot locations, height, and the blowing coefficient will all be parametrically changed.
The resulting flow characteristics will be stored (in the form of pressure fields, vorticity fields,
force coefficients for comparisons with the experimental studies, and for the evaluation of the
noise characteristics of the wing.
-0.2
-0.1
0
0.1
0.2
00.2 0.4 0.6 0.8 1
X/C
Y/C
Upper Surface
Low Surface
Figure 4. The coordinates of the airfoil currently being tested
Appendix B
B - 1
APPENDIX B
A DRAFT COPY OF A DOCTORAL THESIS PROPOSAL
ON THE TOPIC OF THIS REPORT
This appendix includes a copy of the draft proposal for a doctoral thesis on the topic of the
current report by Mr. Scott Munro, a doctoral student working under the direction of Dr. Krish
Ahuja. An important typical feature of a doctoral proposal is a detailed discussion of the
previous. Another important feature of such dissertations is the documentation of initial findings
that justify continuing the work. Since this dissertation is in support of the current NASA grant,
we feel that it will be worthwhile attaching a copy this proposal herewith for the benefit of
readers interested in further details of the findings of the research work to date, as well as a
detailed discourse of the previous work.
It should be noted that since this overall report is only a performance report and is not a
formal report for wide distribution, we are taking the liberty of including here the draft version of
the thesis proposal. It is thus expected that it may not be at par with NASA’s editorial
guidelines. Any feedback provided by the readers will be most welcome as it will only improve
the quality of our research and will add value to the return on investment of NASA funds.
Appendix B
B - 2
Noise Reduction Through Circulation Control Technology
A Thesis Proposal
Submitted to
The Faculty of the
School of Aerospace Engineering
Georgia Institute of Technology
By
Scott E. Munro
In partial Fulfillment
Of the Requirements for the Degree
Doctor of Philosophy
Advisor
Dr. K.K. Ahuja
March 2000
Appendix B
B - 3
TABLE OF CONTENTS
1.0 Introduction B-3
1.1 Motivation B-3
1.2 Objective B-3
1.3 List of Symbols and abbreviations B-4
2.0 Summary of Previous Work B-6
2.1 Circulation Control Wings B-6
2.1.1 The Circulation Control Wing Concept B-6
2.1.2 Circulation Control Wing Development and Applications B-6
2.2 Jet Mixing Noise and Prediction Schemes B-9
3.0 Facilities and Instrumentation (Experimental Set-up) B-12
3.1 Anechoic Flight Simulation Facility and Instrumentation B-12
3.2 Anechoic Jet Noise Facility and Instrumentation B-14
3.2.1 High Aspect Ratio Nozzle (HARN) B-15
3.3 Model Test Facility and Instrumentation B-16
4.0 Technical Approach B-17
4.1 Summary of Proposed Work B-17
4.2 Acoustic Studies B-19
4.2.1 Acoustic Optimization of CCW State of the Art Configurations B-19
4.2.2 High Aspect Ratio Nozzle Measurements B-20
4.2.3 Effect of Jet Turning on Farfield Noise Propagation B-21
4.3 Aerodynamic Verification of Acoustically Optimized CCW B-21
4.4 Initial Results B-22
4.4.1 Acoustic Measurements of Existing CCW configurations B-22
4.4.2 Determining the “Equal Lift” Configuration B-24
4.4.3 Acoustic Measurements of Conventional Wing Configurations B-25
5.0 References B-28
Appendix B
B - 4
1.0 INTRODUCTION
1.1 Motivation
In recent years there has been a significant increase in air travel. This includes, air cargo, passenger travel,
business and general aviation. This increase has caused much greater approach and departure frequency at airports.
This increased business to airports has many effects on the community such as increased commerce and industry.
However, there are drawbacks to the advances in aviation that have allowed this increase. Environmental issues
surrounding airports and air travel have come to the forefront. One of the major environmental dilemmas facing
today’s aircraft industry is the noise pollution from aircraft, especially around the airport. Thus there is a large
emphasis on minimizing the community noise from operation of aircraft in and around the vacinity of the airport.
The FAA already spends millions of dollars a year buying out homes, or making acoustic improvements to homes in
accordance with FFA regulations that require these activities in areas that have increased community noise due to
the airport. Thus airlines, aircraft manufacturers, NASA and the FAA have made reducing aircraft noise a priority.
NASA has proposed a goal of lowering total aircraft noise emissions by 20 dB by year 2020.
In order to meet this goal, NASA and other organizations have been encouraging innovative ideas to help
reduce aircraft noise. Since a major contributor to aircraft noise on approach is airframe noise, reducing this noise
would be helpful in reaching the industry goals. Also, as engine noise is reduced through improved absorbers and
better understanding of jet noise, airframe noise may also become a major contributor on take off as well. The
major airframe noise contributors are the landing gear and flaps that protrude into the flow. Much work has been
done in both of these areas in an effort to reduce their noise emissions. The best solutions would be to have an
aircraft without these protrusions into the flow field. Obviously an aircraft without landing gear would have serious
drawbacks, but there is an alternative system to wing flaps that has shown great promise in maintaining and even
surpassing the lifting benefits of conventional flaps. Circulation control wings (CCW) have been researched and
developed extensively, primarily for the purpose of reducing or replacing the conventional flap system of an aircraft.
CCW was first researched as a short take-off and landing (STOL) device for Navy aircraft. However, over the years
the CCW systems have gone through many configuration designs for many different applications, including versions
for rotorcraft, fighter aircraft, and short haul transports. Even preliminary studies were performed to examine the
possible benefits for a typical transport aircraft such as a Boeing 737 [1]. However, there has been no research
conducted investigating the possible acoustic benefits provided by such a system, other than occasional references to
smaller footprints due to shorter take-off and landing distances.
Since CCW systems have already been shown as an adequate replacement for conventional flap systems in
the aerodynamic realm, they are immediately a candidate for reducing airframe noise since they eliminate the flap
system. An additional advantage may be derived from the fact that the very small heights of the slot nozzle are
likely to produce jet-mixing noise in the ultrasonic frequency range that will be inaudible to humans. However,
there are many issues that need to be resolved before the claims of lower noise can be validated. Since the CCW
system has never been evaluated on an acoustic basis, it must be optimized for this while maintaining, at a
minimum, the lift characteristics of a conventional system. Thus, the proper combination of CCW configuration that
minimizes noise and maximizes the lift must be found. The acoustic effects of many parameters must be
investigated, such as the slot height, slot velocity, and CCW geometric configuration. In order to correctly define
Appendix B
B - 5
the best combination, new areas of research will have to be investigated, including high aspect ratio jet noise, and
the effects of jet turning on its noise propagation. These many issues are the motivation of the present study.
1.2 Objective
The main objective of the present study is to optimize a CCW configuration for minimum noise
characteristics while at a minimum providing the same lifting capability of a conventional flap system. This will be
done by comparing several different CCW configurations to a typical conventional wing system, aerodynamically,
and acoustically. In addition an effort will be made to understand the fundamental characteristics of the noise
sources of a CCW system. For instance, the slot is essentially a high aspect ratio jet, but little research has been
performed on high aspect ratio jet noise. Also, CCW systems turn the slot flow significantly, and this will affect the
propagation of the slot noise into the farfield. Thus, independent studies specifically investigating high aspect ratio
jets and jet turning will be done to determine each independent contribution to the overall noise of a CCW sytem.
The results of this effort will provide basic information to future designers that will help them develop a
CCW system that still maintains the aerodynamic advantages of the CCW concept, but also will have lower noise
emissions than a conventional flap system designed for the same lift capability. There will also be a resulting
increase in the general knowledge of high aspect ratio jets and jet turning effects. This will not only benefit future
CCW studies but will also provide general insights in other applications where these phenomena might occur.
1.3 List of Symbols and Abbreviations
a -- speed of sound.
C -- Chord.
Cl -- Airfoil lift coefficient.
CCW -- Circulation control wing.
CCR -- Circulation control rotor.
Cµ -- qC
Vmj
•
D -- Diameter of round jet exit.
h -- Slot hieght.
I -- Sound intensity.
Mc -- Convection Mach number.
•
m -- mass flow
P -- Sound power.
p -- pressure.
q -- ½ ρV2 (dynamic pressure).
R -- Radial distance from jet exit to measurement location.
r -- radius of CCW surface.
Re -- Reynolds number.
SPL -- Sound Pressure Level.
Appendix B
B - 6
T -- Temperature.
α -- angle of attack
V -- Velocity (different subscripts denote different velocities).
Θ -- Angle of measurement with respect to the flow axis.
ρ -- density.
Subscripts
s -- associated with slot.
T --associated with tunnel freestream.
j -- associated with jet (used in cases where round jets are implied).
o -- ambient condition.
Appendix B
B - 7
2.0 SUMMARY OF PREVIOUS WORK
2.1 Circulation Control Wings
2.2.1 The circulation control wing concept.
The circulation control wing (CCW) concept has been researched since the 1960s It is related to the jet
flap but has some distinct differences that make it more advantageous in many cases. The CCW uses a circular
rather than sharp trailing edge (figure 1). Air is blown tangentially along the upper surface from a plenum supply
inside the wing through a slot just upstream of the rounded trailing edge. Blowing under the right conditions moves
the upper surface separation point around the trailing edge, thus changing the trailing edge stagnation point location
and hence the circulation for the entire wing. The high speed air moving along the surface also causes a suction
peak in this region and contributes to increased lift.
The slot flow remains attached to the surface due to the so-called Coanda effect. This effect is produced by
a balance within the jet sheet between the centrifugal force and the low static pressure generated by the high velocity
slot flow. At low blowing velocities, the tangential blowing behaves similar to a boundary layer control device by
adding energy to the slow moving flow near the surface. At higher blowing rates, the lift is increased by the change
in circulation described above. As mentioned, the concept appears similar to a blown flap configuration.
However, the rounded trailing edge makes for some rather significant differences. With blown flaps, the
trailing edge is still sharp, resulting in a fixed trailing edge stagnation point. The blown flap is also much larger than
the trailing edge surface employed in a CCW configuration. This results in much higher mass flow requirements for
the slot to achieve higher lift. As such with blown flap configurations it is not possible to modify the circulation by
controlling the trailing edge stagnation point. The blown flap usually requires a rather significant sized flap that has
to be mechanically deployed and retracted, whereas the CCW can be designed without any mechanical movements.
More recent developments of CCW technology have yielded a device very similar to a blown flap. Rather than a
flat blowing surface in the case of a blown flap, the CCW flap system has a curved top surface, essentially similar to
a slice of a circular edge (figure 2). The flap allows for a low drag cruise configuration, but still provides much
more turning than a blown flap because of the curved upper surface.
2.1.2 Circulation Control Wing Development and Applications
There are many potential uses for circulation control. However, the two applications that have received the
most research attention have been circulation control rotors (CCR) and CCW applied to an aircraft for short take-off
and landing (STOL) capability. Much of this work was done for the Navy as development work for carrier based
aircraft. Much of this work is documented in reference [2]. This reference is a bibliography of abstracts of nearly
all early work in CCW and CCR work.
Dunham developed a theoretical basis for circulation control and applied it to a circular cylinder with
tangential blowing on the upper surface [3] and review by Henderson in reference [4]. With minimal blowing, the
boundary layer is re-energized, delaying separation on the upper surface. As mentioned above, in the case of an
airfoil with a sharp trailing edge, the trailing edge stagnation point is fixed by the Kutta-Joukowski condition.
Appendix B
B - 8
Replacing the sharp trailing edge by a cylinder changes the trailing edge condition. The new requirement is that
upper and lower separation occurs at the same pressure [4]. Thus, if the upper surface pressure rise is delayed, upper
surface separation is delayed, and hence the stagnation point moves further around to the lower surface increasing
the circulation. This ability to move the stagnation point is maximized by choosing a circular cylinder [4].
Dunham's theoretical method essentially calculated the boundary layer around the cylinder until the separation
criteria was met. Dunham was able to incorporate the tangential jet into the calculations and arrived at the result (for
a circular cylinder)
1
1
~+
∞
c
c
r
r
h
V
Vs
lKcδ
γβ
α
Noting that δ/c is the boundary layer thickness at the slot, and is
5/1
Re
K
c=
δ
Even though α, β, and γ are constants that are undetermined, one can see the major factors that will enhance lift are
high jet velocity flow out of the slot, large slot heights, and a large trailing edge radius. However, as with any
design issue there are limits. The jet velocity is limited by the practical issue of maximum pressure obtainable in an
aircraft configuration (typically bleed air from an engine) and or physical systems available in ground test facilities.
Larger slot heights are somewhat beneficial; however, too large a slot height will upset the Coanda effect balance
with too thick a jet sheet and the jet will separate prematurely. To give the most efficient turning, large radii are
desired [4, 5].
Earlier designers used large radius trailing edges to maximize the lift benefit. However, these designs
suffered from higher drag [4, 5, 6]. Reducing the trailing edge radius will reduce the overall drag, but the high lift
will also be reduced, and the smaller radius is also much more susceptible to separation around the sharper turn [5].
Thus a compromise was needed. The large circular trailing edges evolved into a circular hinged flap [1, 5, and 7].
The hinged flap was a compromise of several desired features. The flap had a curved upper surface, like the
cylindrical trailing edge, but a flat lower surface. The curved upper surface and flap deflection gave some flow
turning ability, but the flat bottom surface and the sharp trailing edge virtually returned the airfoil to its clean shape
when the flap was retracted. This produces a major drag reductions compared with any of the cylindrical trailing
edge configurations. The flap does require some mechanical function, which increases the weight, but allows some
mechanical camber adjustment should the circulation control system fail. Overall, the hinged flap design still
maintained most of the circulation control lift advantages but greatly reduced the drag problem associated with the
circular trailing edge system.
This flap is similar to a blown flap, however, it is important to note that the upper surface is curved. The
curvature is either a curve built from a single radius, or multiple radii, see figure 2 (the dimensions shown here are
for the test configuration that was employed in the present experimental and computational study). This produces a
partial CCW surface but maintains low drag in cruise configuration. Once the flap is deflected, the blown flow
remains attached and leaves the trailing edge at the angle of the surface. Thus the trailing edge stagnation point is
fixed. However the circulation can be controlled by combination of blowing amount and flap deflection. Some lift
Appendix B
B - 9
can be produced without any flap deflection, thus a low drag (profile drag) high lift configuration is possible, but
with large flap deflection and high blowing the lift and drag are significantly increased. The flap itself has several
mechanical advantages compared to conventional Fowler flap systems. The flap is about 1/4 to 1/3 the size of a
conventional flap. This means lower flap weight, and thus fewer structural components are required to hold it in
place [7]. The flap is also a simple hinged flap, rather than a complex Fowler type flap that requires complex
gearing and tracks, which most likely contribute to airframe noise just on their own. The reduced size and simplicity
of the CCW system even with a small flap clearly offers some advantage over a conventional system even if one just
considers them to have equal lifting capability. For these reasons the CCW flap system has been the main CCW
concept studied in recent years.
There are many applications for this technology. As discussed above, the initial intention for development
of the system was for short take-off and landing (STOL) capability. The interest was primarily from the Navy,
looking for ways to improve aircraft operation from carriers. The Navy has even gone so far as to sponsor a full-
scale flight test program on an A-6 Intruder. The design, tests and results are documented in references [2, 6, 8, and
9]. This test used a large cylindrical trailing edge. The experimenters were aware of the drag penalty but were
mainly trying to prove the concept at full scale. Since that full-scale test, considerable additional design work
incorporating the smaller cylinders and the hinged flaps has been performed.
In fact, there was interest designing a system for a transport aircraft. Research was done to investigate
applying the circulation control system to a Boeing 737 aircraft. This research included experiments and
computational work looking at a design that would use trailing edge blowing to eliminate the trailing edge flaps, and
leading edge blowing to eliminate the need for leading edge slats. A summary of the effort is documented in
reference [1].
It should be noted that the computational and experimental work used the same wing design used in the
current study. The wing currently used has a leading edge and trailing edge blowing slots (independent) and the
hinged flap configuration described above. The study showed that a large-scale CCW system could be used on a
737 type aircraft and significantly improve the STOL capabilities.
However, to our knowledge, the only noise study of a configuration involving CCW is that by Salikuddin,
Brown and Ahuja [20]. In their study, they examined the changes in noise produced by an upper surface blowing
with and without CCW. No one has however, compared the acoustic signature of a CCW and a conventional wing
for the same lift performance. This study aims to do just that. It could potentially improve the lifting ability of the
aircraft while at the same time reduce the overall noise associated with the wing.
Since the system developed in reference [1] is directly applicable to the airlines, a large market for reduced
airframe noise, it was chosen as an appropriate starting place for this study. Thus, the test model wing used in
reference [1] is being used as the test model for this study. The reasoning behind using this wing for the current
study was two fold, 1) the wing was available and already fabricated and 2) the design is the latest step in the
evolution of CCW systems. There is no reason to perform acoustic analysis on a CCW system that is not practical
aerodynamically.
It should be noted that there are other potential uses for circulation control, including automotive
applications [24] and helicopters. There has been a significant amount of research on use of circulation control in
helicopters, where the main focus of this work was to eliminate the complex mechanical structure needed to produce
Appendix B
B - 10
cyclic and collective pitch for the rotor. References [2, 10, and 11] provide examples of the research done in this
area. However, since the current work is directed at aircraft with flap systems, only this brief mention of circulation
control on helicopters will be done. For further information, it is suggested that the reader begin with the mentioned
references.
2.2 Jet Mixing Noise and Prediction Schemes
As soon as jet and rocket engines began making their way into the aircraft designs, the noise from these
new types of engines became an issue. In some cases it was more for controlling damage, such as in the case of a
rocket launch, where the launch area is subjected to a large amount of noise from the rocket motors during the
launch. The other major issue came with increased jet travel and jet aircraft activity around airports. The new jet
aircraft were much louder and more annoying to the surrounding population.
Thus, research into jet noise soon began to emerge. Much of the early theoretical gains in jet noise prediction came
from Lighthill's work on round jets. Various versions of this work are found in references [12, 13, 14, and 15].
Lighthill derived equations for a turbulent flow including Reynolds stresses and the associated turbulent terms. He
also showed that the dominant stress term was ρvivj at low Mach numbers. Since turbulent fluctuations correlate
well for points within a volume on the scale of the typical eddy size, Lighthill proposed that acoustic sources
associated with the turbulent fluctuations at these points will be coherent. Thus, the distribution of quadrupole
sources in the volume will radiate sound similar to a single quadrupole equal in strength to the combined
distribution. Essentially the noise associated with a particular eddy is represented by a quadrupole source. Lighthill
went on to assume that the jet flow was made up of a number of eddies, and thus a similar number of these point
quadrupoles representing eddies. From this physical model, several relationships were derived, including
'Lighthill's eighth power law' relationships for the sound intensity and the sound power
: 2
R
5
o
a
o
2
D
8
j
V
2
m
~Iρ
ρ and 5
o
a
o
2
D
8
j
V
2
m
~Pρ
ρ.
The derivation shown above has mostly been applied to round jets of diameter D. Secondly, there are important
relations that are shown, specifically that the sound intensity is proportional to the eighth power of velocity and
inversely proportional to the square of the radius between the source and observer. The final note on these
equations is that they also assume that the eddies (or quadrupoles) convected at a very low Mach number. When
the eddies are convected at a higher Mach number (Mc) the analysis must take this into account by shifting to the
reference frame of the convecting eddy. When backed out to the observer, the equations change to
5
c
2
R
5
o
a
o
2
D
8
j
V
2
m))cos(M1(~I−
ρ
ρΘ− and 42
c
2
c
5
o
a
o
2
D
8
j
V
2
m
)M1(
)M1(
~P−
+
ρ
ρ.
In the intensity equation, Θ is the angle off the jet axis. These equations predict the magnitude of the of acoustic
noise, but say nothing of frequency. However, one can again go back to the eddy which is essentially the driving
force behind the noise. Near the exit of the nozzle where the mixing region is small, the turbulence is dominated by
small eddies, thus higher frequency noise is associated with the small length scale. As the shear layer grows, the
larger eddies further downstream are believed to be responsible for lower frequency jet noise (Ahuja, 1973). But
notice that these characteristics are dependent on the geometry and mixing characteristics of the jet. Thus, the
Appendix B
B - 11
frequencies must also scale in order to be able to predict the entire spectrum of jet noise. The frequency scaling is
taken into account by non-dimensionalizing the frequency into a Strouhul number and accounting for the moving
sources. This relation is
)cos1(Θ−c
V
fD M
Most of Lighthill's theory has been experimentally verified for round jets. One key study in this area was
performed by Ahuja, and Ahuja and Bushell [16, 17]. They made careful measurements of jet noise for 3 different
diameter round jets. An effort was made to eliminate all other possible sources of noise, such as valve and flow
noise from upstream in the nozzle system. Ahuja verified the data by scaling all his data to the same condition,
which would collapse all the data if Lighthill’s theory was correct. This is a fairly straightforward process when
working with sound pressure levels. Since the sound pressure level (SPL) is defined as
=ref
I
I
SPL log10
where Iref
= 10-12 W/m2. Thus, if the intensity is normalized by some “standard” intensity and a “standard” SPL is
solved for, the result is
( )
5
228
cos1log10log10log10log10"tan"−
Θ−−
+
−
−= c
sss M
R
R
D
D
V
V
SPLSPLdards
where the variables with an ‘s’ subscript signify conditions of the “standard” case, i.e. Vs = 100 ft/s, Ds = 1”, Rs =
10 ft. Thus, any SPL measurement from a jet could be transformed, or scaled to the SPL for this standard case. If
Lighthill’s theory holds, then jet noise from experiments could all be collapsed into one curve if the by plotting the
“standard” SPL for all data versus the normalized frequency. This is very powerful information since the reverse
could be done as well. Any SPL data could be scaled from one geometry, distance, or velocity to predict what the
noise would be at another case. Ahuja's experimental data for unheated jets agreed with many of Lighthill's
predictions and Lighthill's theories are still the standard for predicting jet noise and extrapolating noise from one
system to another by using the velocity, diameter, and radius factors derived in the above equations. However, one
must keep in mind that all of this work was done for a round jet so it does have some limits.
Since this work, there has been a significant amount of work on jet noise, from acoustically exciting the jet,
to mixing studies, to co-axial flows (recent work of Dr. Tam at Florida State Universtiy is particularly noteworthy).
However, again, most of this work has been done with round jets. There has been some work in rectangular jets.
Some have tried to formulate an equivalent diameter so the round jet predictions can be used. But the aspect ratios
considered 'high' are two orders of magnitude smaller than the typical aspect ratios of a CCW system, so there is still
a question as to how to scale the geometry, if the other portions of the theory hold in this case.
Appendix B
B - 12
3.0 FACILITIES AND INSTRUMENTATION
There are 3 facilities that are possibly going to be utilized in the present experiments. These include the
anechoic flight simulation facility (AFSF), the anechoic static jet facility (ASJF), and the Model Test Facility
(MTF). All are GTRI facilities located at the Cobb County Research Facility.
The AFSF and ASJF are primarily acoustic test facilities and are the primary facilities used thus far. The
AFSF operates in an open jet wind tunnel configuration, but there is no balance system for direct force and moment
measurement. Critical pressures are monitored, but the AFSF is not set up to measure a large amount of pressures,
i.e., similar to what might be expected for a reasonable pressure distribution. The ASJF is an anechoic chamber with
a plenum air source and a collector which allow for installation of nozzle configurations for acoustic measurements.
There is no external flow, all measurements are static jet measurements. The MTF is strictly an aerodynamic
facility and has no capability to make acoustic measurements. It, however, does have a balance system and
instrumentation which allows for many pressures to be monitored and recorded.
3.1 Anechoic Flight Simulation Facility
The AFSF is a semi anechoic facility that allows acoustic measurements to be made in the presence of a
freestream (see figure 3). The tunnel has a rectangular inlet which converges down to a 28 in. round duct. The duct
terminates in an anechoic room as an open jet. Protruding out from the downstream wall is the collector which is 4
ft. wide by 5 ft. high. The collector duct extends outside the building and ends at a centrifugal fan powered by a
diesel engine. The facility is open circuit, drawing air from outdoors. Acoustic foam wedges are mounted on all the
walls of the anechoic room. The open jet duct, and the collector are also covered with foam to prevent reflections.
The room itself is isolated from the rest of the building by gap between the room wall and the wall for the rest of the
building. The walls of the collector are also filled with sand to prevent noise from outdoors from propagating into
the collector and up into the anechoic room.
In the current experiments, the wings are mounted via mounting brackets to a support for the open jet. This
locates the wing across the jet opening immediately downstream of the end of the duct. Figure 4 shows one of the
conventional wings mounted at the exit of the open jet. Although the wing models have a large number of pressure
ports that would enable one to obtain a pressure distribution. They were not measured in this facility as the main
purpose of using this facility was to acquire acoustic data. Detailed pressures are however, obtained in the MTF
described below. Critical pressures such as the ambient pressure in the chamber, the plenum pressure for the slot,
pressures in the air supply line venturi mass flow meter, and pressure in the inlet (for freestream velocity) are
monitored on individual pressure transducers and manually recorded for each test point. There is real-time software
located on a computer in the control room which monitors the pressures in the mass flow meter and continually
displays a mass flow through the supply line.
Acoustic measurements are made with B & K, 4135, 1/4 in. microphones. One microphone is mounted on
a traverse system that can translate the microphone from angles of 30o to 90o (where 0o is the freestream direction).
Appendix B
B - 13
This system is set such that all measurements are made in the fly-over plane. The microphone is connected to a
multi channel digital frequency analyzer which is run by software on a PC located in the control room. Once
calibrated, the microphone data can be recorded at the push of a button and saved to an output file containing
frequency spectra for all the channels of interest. The software also allows such changes as number of averages,
frequency span, and frequency resolution (∆f). The output files can be plotted using virtually any plotting program.
In addition to acoustic measurements, some critical pressures must be measured to monitor tunnel
conditions. The main measurement that must be made is tunnel velocity. The tunnel has a well established
calibration for tunnel velocity based on two pressures measured, chamber ambient pressure and a pressure measured
at a specific location in the inlet. The velocity equation is
where ∞
V is in ft/s, Pwall is the inlet pressure mentioned above (gauge pressure), Pchamber is the ambient pressure in
the chamber (absolute pressure), and T is the ambient temperature in Rankine. The exact history of this equation is
unknown by the current researcher, however it has been used for years in this facility. It is suspected that the
constant is a combination of physical constants, calibration constants and unit conversions. Before testing, this
velocity equation was verified by inserting a pitot probe at the exit of the open jet. The velocity calculated from the
equation and the pitot probe matched very well (see figure 5). Thus, the velocity equation was used to calculate
velocity for tests since it is non-intrusive into the freestream.
The final piece of equipment used in the AFSF is the blowing system for the CCW wing. Since this system
can be transported to any of the facilities it will be described as an independent component here. Figure 6 shows a
schematic of the blowing system. The system simply consists of high-pressure 3/4 in. tubing, a mass flow venturi,
pressure gauges and a muffler. On the upstream end, the tubing is connected to an existing high pressure line with a
control valve upstream (A connection to the high pressure air system is available in most of the facilities at GTRI).
Next in line is the mass flow venturi. Once the flow passes through the venturi, it goes through more tubing to an
in-house built muffler. Downstream of the muffler, the air passes through more tubing to inlets for the CCW
plenum.
The venturi mass flow meter works by monitoring pressures at the inlet and throat of a convergent-
divergent nozzle. The flow is usually assumed isentropic and calorically perfect, and the mass flow is then
calculated by using the governing equations. There are two basic approaches that are taken, either the flow is
assumed 1-dimensional and incompressible, or the flow is assumed quasi-1dimensional through a variable area duct.
In either case, the accuracy and efficiency suffers greatly if the flow in the throat is transonic or greater. Based on
the upstream conditions and the area ratio, the Mach number in the throat can be calculated. From this, the
isentropic relations can be used to find the local density and speed of sound. The mass flow is calculated using these
parameters. Most venturi meters are calibrated by the manufacturer and were shipped with an empirical correction
curve based usually on the Reynolds number of the flow through the meter. This correction factor accounts for
losses associated with skin friction on the walls, separation and other losses.
T8356.169Vchamber
Pwall
P
=
∞
Appendix B
B - 14
By monitoring the pressures in the inlet and at the throat, the mass flow through the meter can be
determined. In the present experiments, the pressures are measured using two Synsem pressure transducers, one 0 -
100 psig gauge, and one 0 - 30 psig gauge. The 0 -100 gauge is used to sense the inlet pressure. Due to losses in the
flow system, and since the mass flow meter is upstream of the muffler, pressures in the meter will be relatively high
compared to the wing plenum pressures. Thus far, for a plenum pressure of 25 psig, upstream pressures at the mass
flow meter have been around 60 psig. The 0 -30 psig gauge is used to sense the differential pressure between the
inlet and throat pressures. This is done rather than using another 0 -100 psig gauge because much greater accuracy
can be achieved. It is found that velocity changes through the meter is rather small due to the very high density
associated with the 60 psi total pressure at the meter's location. Even though near the plenum the flow might be
moving at 300 ft/s through the supply line with a total pressure of 25 psig, upstream with the total pressure of 60
psig much lower velocities are required to maintain the same mass flow. Thus, a much more sensitive pressure
transducer is used to accurately sense the pressure differential between the inlet and throat of the meter. Both of the
pressures are fed into an in-house Labview code that continually displays the mass flow. The code incorporates the
manufacturer corrections provided with the venturi.
In the above paragraphs the techniques for operation and acquiring data were discussed. However there is
still one important issue that must be resolved before any acoustic data can be said to truly represent flight data. The
free flight test facility introduces a shear layer between the test model and the observer that is not present in the real
flight case. The shear layer is caused by the mean flow exiting into the anechoic room. This could be avoided by
placing the microphones in the flow, but then there are other issues that need to be dealt with, such as the flow
impinging on the microphone and causing flow noise not associated with the model. Also, unless the tunnel is large
there could be a problem with mounting microphones far enough away from the model such that they are in the
acoustic far field.
However, in this case, when the tunnel was designed and built, the open jet shear layer was chosen as the
problem to overcome. Ahuja et. al performed a study in this same facility and developed techniques to properly
account for effects of the shear layer on the propagation and magnitude sound from the model through the shear
layer. This is documented in reference [18]. This includes corrections for emission angle, magnitude, and
scattering due to the shear layer. These corrections will be applied to data when needed.
3.2 Anechoic Static Jet Facility
The anechoic static jet facility is also located at GTRI at the CCRF. Figure 7 shows a schematic of the
ASJF. The facility is based around a 10 in. acoustically treated plenum chamber that contracts to a 4 in. exit. There
is also a secondary plenum over the primary plenum that exits through an outer ring of exit ports on the plenum face.
This configuration allows for two independent coaxial flows. The flows can be independently controlled and
heated. Up to the two plenum chambers, the two flow systems are identical but independent. Both have control
valves that allow the pressure in the plenum to be regulated manually by the user. Both have a propane burner
system in the flow line that can heat the flow to about 1200 F when desired.
Appendix B
B - 15
The plenums dump into a large anechoic chamber with a collector on the opposite wall so the jet flow can
escape. Since only the primary plenum will be used in the current experiments, only it will be described further.
The primary plenum contracts to a 4 in. exit flanged face. Any nozzle fitted with the standard flanged fitting can be
bolted to the plenum face. For the current experiments a high aspect ratio nozzle (HARN) was designed, build and
attached to the plenum face (see figure 8 for a drawing of the HARN). More details on the HARN will be provided
in the next section. The nozzle extends 38 in. beyond the plenum face. The exit of the HARN is the reference point
for a microphone arc located on a 10 ft. radius from the center of the HARN exit. Microphones are locatedin a polar
arc passing through the jet axis at angles of 30, 40, 50, 60, 70, 80, 90, 100 and 110 degrees with respect to the jet
flow direction (Θ). The nozzle can be mounted to the plenum at several azimuthal positions (Φ), allowing the arc to
be in the lateral or fly-over plane by only changing the orientation of the nozzle. The microphones are connected to
extension cables, which are connected to the same frequency analyzer as used in the AFSF. The pressure and
temperature in the plenum of the facility are monitored visually via a pitot probe and a thermocouple connected to
visual displays in the facility control room. It is assumed that the pressures and temperatures in the plenum are the
same as those in the plenum of the HARN nozzle. These are reasonable assumptions since the flow is unheated for
these tests and since the area ratio for the 8 square inch inlet to the outlet of the HARN is to range from a minimum
of about 1.5 to a maximum of 100, keeping in mind that the current experiments are in the range from about 8 to
100. The temperature should be relatively constant through the system, and the high contraction ratio in the HARN
should allow its contraction area to be plenum like, with very little loss between it and the main facility plenum.
The plenum and atmospheric conditions are used to determine the isentropic jet conditions for the nozzle.
3.2.1 High Aspect Ratio Nozzle HARN
The HARN was fabricated at GTRI in order to perform jet noise studies on very high aspect ratio jet flows.
The primary focus was to provide a similar jet flow model to the CCW wing, in scale and flow, separate from the
other aerodynamic influences incurred in the CCW system. Because of the interest in maintaining similarity
between the HARN and the CCW being tested, the HARN was designed to be 30 in. wide (see figure 8). The
HARN mounts to a round to rectangular transition duct section (2.75 in. square exit) that mounts directly to the 4 in.
opening on the plenum. It gradually contracts to the desired slot height in one dimension and expands to the 30 in.
width in the other. Both of these dimension changes occur over about 30 inches of axial length. The HARN was
designed with exit plates that are adjustable. This allows the slot to have potentially an infinite number of slot
heights within the bounds of fully closed to 0.25 in. For the current study, most of the heights will be similar to
those in the CCW tests, ranging from 0.003 in. to 0.020 in. However, if it is felt that more slot heights will further
our understanding more heights will be added to the test matrix.
The HARN was fabricated out of aluminum with an intended maximum pressure of about 25 psig, enough
to produce a slot exit Mach number slightly over 1.2. The outside surface of the HARN is sloped towards the exit,
so that the entrained flow is more nozzle-like rather than a wall-jet. A curved attachment piece is currently being
Appendix B
B - 16
fabricated, again maintaining the shape and size of the CCW system. It will be attached to the HARN via a piece
that will allow it to change angles similar to deflecting the flap on the CCW.
3.3 Model Test Facility
The model test facility at GTRI has been used for several CCW experiments in the past. The test section is
30" x 43". The CCW is mounted on a balance such that it spans the 30 in. width of the tunnel. Since the wingspan
is 30 in. the installed wing is essentially a 2-D airfoil when mounted in the tunnel. The conventional wings are
mounted in the same manner. This is similar to the set-up in the AFSF since the wings span the entire open jet flow.
The floor and ceiling of the MTF are fitted with blowing slots originally designed for ground effect studies.
However, use of blowing in 2-D wing studies helps reduce wing/wall junction effects by eliminating the momentum
deficit near the tunnel walls and will be used for the present study while making the aerodynamic-performance
measurements. The CCW and conventional wings will be mounted on a 6 degree of freedom balance system. In
addition pressure distributions over the wings can be measured using scanivalve pressure measurement systems.
The tunnel can achieve similar freestream velocities to the AFSF and is a closed-circuit continuous run system.
Appendix B
B - 17
4.0 TECHNICAL APPROACH
4.1 Summary of Proposed work
As mentioned in previous sections, much work has been done in the area of circulation control. However,
most, if not all of this work has focused on the aerodynamic advantages of circulation control wings and possible
applications based on aerodynamic performance. Little has been done to exploit the possible acoustic advantages of
a CCW system. Thus, the proposed work will focus on optimizing a CCW system for low noise impact while
maintaining aerodynamic performance sufficient for use in a similar situation as a conventional flapped wing
configuration. This must be done in two parts, the acoustic and the aerodynamic. It should be mentioned that this
study is primarily an experimental aeroacoustic investigation of the CCW system. In conjunction with this work,
there is also computational work being performed under the same grant. However, this study will focus on
experimental work and only use computational work to compare with findings of the experimental tests.
The first step is to determine if and how a CCW configuration can have lower noise than a conventional
system. This step involves side by side comparison of representative configurations under the same conditions, i.e.,
the same mean flow and lift conditions. Since there are several variations of CCW systems that have been
researched, a basic study of different CCW configurations was done. Since much work on CCW systems has been
done at GTRI, a CCW model used in previous studies was used. This also allowed for comparison with previous
aerodynamic data on the same model, thus making the current work an extension of the current database on this
model. This model had the same baseline airfoil shape (a Lockheed supercritical airfoil LG1616), but had many
different detatchable CCW trailing edge configurations. Several blowing configurations (blowing slot height, Cµ)
were compared to a conventional wing system. The conventional wing system also had the LG1616 supercritical
airfoil shape for most of the airfoil. However, its trailing edge was altered with a cut-out for a stowed flap. For this
model, the cut-out is open, and a single slotted Fowler flap was attached. The flap was deflected 40 degrees off the
chord line to simulate a landing configuration. Initially the two systems were compared at the same angle of attack
to keep the number of variable parameters reasonable. But since the CCW system can increase lift without changing
angle of attack it may have another advantage to compare different angles of attack that produce the same lift.
Initial acoustic evaluation was performed in the flight simulation facility at GTRI. One microphone located
in the far field was used to acquire the noise spectra of each configuration. The use of only one microphone at this
stage simplified the data acquisition process significantly and allowed for a much broader study of different
configurations in the same amount of time. This was enough to get a basic idea of the noise characteristics of
different configurations. Once basic noise spectra of the CCW and conventional wing configurations are acquired at
several mean flow velocities and angle of attack, specific cases where the different configurations with the same lift
coefficient were compared directly. The lift data was gathered from previous studies for this initial comparison.
Configurations (combinations of angle of attack, slot height, and Cµ) that had similar or lower noise than
the conventional system for the same lift were then considered for more in-depth noise and aerodynamic
Appendix B
B - 18
measurements. All configurations that produce acoustic advantages will be aerodynamically evaluated. The
models will be placed in the MTF wind tunnel at GTRI. Force and pressure measurements will be made at all
conditions of interest. This will provide a database of “equal lift” configurations that can then be evaluated
acoustically in depth. These tests are critical since initial data has already shown trends that indicate the best
acoustic configurations have not been tested aerodynamically. Essentially the aerodynamic study will provide a test
matrix of different combinations of angle of attack, slot height, and Cµ that will all have equal aerodynamic
qualities. The most important of these will be to maintain equal lift; however, drag and pitching moment will also
be monitored. The chosen configurations will be placed back in the flight simulation facility at GTRI. A
microphone will be placed at different angles to the trailing edge of the wing in the flyover plane by using the
traverse system described previousely. Data will be acquired at a variety of freestream and angle-of-attack
conditions, in addition to several blowing configurations that gave positive results in the case of the CC wings.
This investigation will give much insight into the noise characteristics of a CCW wing system and how it
compares with a conventional flap configuration. However, identifying individual noise sources in the system will
be difficult with the entire CCW system. Much value will be added if individual noise components can be studied
and used to determine how much each part affects the noise of the entire system. Possible individual contributors
are the jet flow from the slot, the changed directivity due to curvature, or perhaps the jet impinging on freestream
flow. Although much work has been done in the area of jets, no one has extensively looked at the properties of
extremely high aspect-ratio jets, similar to the CCW slot. In the current study, aspect ratios range from 1,200 to
10,000. There has been acoustic work on rectangular jets, but the aspect ratios have rarely been greater than 10.
Thus, to study the unique characteristics of a CCW slot-like aspect-ratio jets, a high aspect ratio nozzle (HARN) has
been fabricated. The HARN will be mounted in the Anechoic Chamber at GTRI where it will be used to study the
effect of aspect ratio on the jet noise. Acoustic measurements will be taken at several jet Mach numbers, slot
heights, and aspect ratios. Acoustic data will again be taken with several microphones located in the acoustic
farfield at several angles relative to the jet flow.
Attachments to the HARN will also allow for investigation of jet turning on the directivity of the noise. A
variety of curved surfaces will be fabricated to simulate the CCW cylindrical surface that creates the jet turning.
Acoustic measurements will be taken in a similar fashion to the aspect ratio tests discussed above.
Results from both of these studies will be compared to jet theory directivity and level predictions to
ascertain whether or not the noise from these unique aspects of the CCW technology can be predicted by existing
theory or by some correction to existing theory. This information could be invaluable to future designers of CCW
systems that have to not only consider the aerodynamic aspects, but have to design an environmentally friendly
system.
Appendix B
B - 19
4.2 Acoustic Studies
4.2.1 Acoustic Optimization of Existing CCW State-of-the-Art Configurations
Since the CCW concept has been around for nearly 40 years there have been many advances, changes, and
modifications to the basic concept to improve its overall performance. To attempt to acoustically test all the
different configurations would be unreasonable, and also poor scientific procedure since many of the changes were
made to improve the system. Thus there is little reason to acoustically test a system that will never be used because
it has been technologically surpassed by a newer version. Thus, the goal of the current study is to investigate two or
three of the best performing CCW configurations. This will provide a database to future designers that will enable
them to use the best aerodynamic and acoustic CCW design to fit their needs.
If one looks at the progression of CCW technology, it started from a rather large trailing edge cylinder.
However, as discussed in previous sections, the large trailing edge created larger than desired cruise drag
performance (assuming the device was not retracted into the wing during cruise). Thus, the use of small cylindrical
trailing edges was preferred. This reduced the drag problem, but did not eliminate it, and in addition, the lift
advantages is reduced since the smaller-radii cylinders were more prone to jet separation that would destroy the
CCW lift. The need to maximize the CCW benefits but reduce the cruise drag led designers down a path similar to a
blown flap device.
Instead of having a complete cylinder along the trailing edge, a large radius partial cylinder was used. The
device was incorporated into a small flap (on the order of 1/3 the size of a conventional flap) that deflected about a
hinge point. This provided a large CCW turning radius, low drag with the flap retracted in cruise, and if the CCW
system failed some mechanical camber adjustment was available for increased lift. However the compromise did
sacrifice the maximum jet turning of the completely cylindrical trailing edge. Maximum turning of the jet was
limited to the angle of the flow coming off the curved flap plus the flap deflection angle. One must keep in mind
that the angle leaving the surface is the angle of the line tangent to the surface. Thus, even without any flap
deflection, the flow leaves the trailing edge at some angle with respect to the lower surface. This results in some
turning of the flow with no flap deflection.
GTRI has an existing CCW with a small flap. The flap can be set at a variety of angles. The flap can also
be removed and other types of CCW trailing edge devices can be installed. A conventional wing configuration with
the same basic airfoil shape is also available to make comparison measurements. These models were chosen as the
models for the optimization study. Based on previous aerodynamic work, the CCW with 90 degree flap deflection
was chosen as the beginning point for the study. This had the best overall aerodynamic design of several
configurations tested. The flap was eventually adjusted to 30 degrees deflection to prevent flap-edge vortex
shedding noise that was measured. It should be noted that the tone was extremely loud (20 dB above the other
noise) without any blowing, but was reduced significantly with minimal blowing when the flap was at 90 degrees
deflection. At 30 degrees deflection, the tone was much lower without blowing and virtually eliminated with very
small amounts (this result will be discussed more thoroughly in section 3.3.1). Three slot heights were chosen for
Appendix B
B - 20
the optimization study; 0.003 in., 0.006 in., and 0.012 in. 0.006 in. and 0.012 in. These dimensions were chosen
because they were typical slot heights used in our earlier aerodynamic studies [1] while 0.003 in. was added when
initial results showed noise reduction for very small slot heights. A wide range of slot Mach numbers was
evaluated, ranging from 0.6 to 1.2. The higher Mach numbers were chosen because high jet velocities have shown
large aerodynamic benefits. The lower limit was set at 0.6 because there was worry that the high Mach numbers
might produce too much jet noise even though they had significant aerodynamic benefit. As it turns out, initial
results indicate that even lower slot-jet velocities may provide sufficient lift while having much better acoustic
benefits. Acoustic performance of all of the CCW test configurations were compared with a conventional flap
configuration. The only difference between the models was the trailing edge containing the retracted flap cut-out
and a conventional flap (with chord about 30% of the wing chord) deflected to 30 or 40 degrees to simulate a
landing configuration.
All configurations were tested in the anechoic flight simulation facility at GTRI. One microphone
was located at 90 degrees to the freestream, below the trailing edge of the wing. Data was acquired for each test
configuration at freestream speeds of 100, 150, 200, and 250 ft/s (nominal) and at geometric angles of attack of 0, 7
and 14 degrees.
4.2.2 High Aspect Ratio Nozzle Measurements
Although the CCW can provide essential data and results for basic analysis of the questions concerning the
noise of a possible CCW flap system, there are many issues that are not clear when just examining the CCW system
itself. This is especially true since the available CCW models were not designed with acoustic concerns in mind.
The internal design of the air supply inlets, plenums and slot may be fine for aerodynamic purposes, but are not
designed to limit the flow noise generated in these areas. Thus there is the possibility of contamination of the
acoustic data by noise generated internally. While it is admitted that this problem will have to be dealt with
eventually in the design of a full scale system, this internal noise is much more specific to the individual system and
not to the general noise characteristics of a CCW type system. This study will focus on the noise generated by flow
once it leaves the slot, not noise due to sharp edges inside the plenum, or other internal noise sources. Thus, there is
a possible problem with using available CCW models for all flow conditions. Unfortunately to build an airfoil with
a “low noise” internal system was cost prohibitive due to the intricate machining required to fabricate an appropriate
model at the scale mandated by the facilities. Thus, a generic test nozzle was built at a much lower cost that could
study the different aspects of the noise generated by a jet of similar dimension to the CCW slot with extremely high
aspect ratio.
As mentioned in section 3.2.1, a high aspect ratio nozzle (HARN) was fabricated for evaluation of very
high aspect-ratio jets. The HARN will be mounted in the anechoic static jet facility at GTRI. The HARN has
adjustable slot heights from fully closed to 0.25 in. With its 30 in. width, the nozzle is capable of aspect ratios from
120 to infinity, in theory. However, its upper limit will be limited to the measurability of extremely small slot
heights. Microphones will be mounted on an arc of constant radius from the nozzle in the acoustic farfield. The
Appendix B
B - 21
nozzle can be oriented at different angles on the mounting plenum, thus the same microphones can be used to
measure in several different planes with respect to the nozzle (i.e., the flyover plane, lateral plane, or some plane in
between).
The slot height and jet Mach numbers will be set at the same heights used in the CCW experiments in the
flight simulation facility. If there is a need for more test points to clarify results they will be added when deemed
necessary. These results will be compared to theoretical jet noise prediction methods that are well established and
currently used to predict noise for round jets.
4.2.3 Effect of Jet Turning on Farfield Noise Propagation
Attachments can be added to the HARN nozzle to give it a cylindrical trailing edge similar to a CCW. A
curved flap like surface has also been fabricated to simulate the turning flap of the CCW. This will enable
independent investigation of two major effects; 1) high aspect ratio jets mentioned above, and 2) turning of the jet
flow by a Coanda surface. This second effect is likely to change the directivity of jet noise compared to the no jet-
turning case. Thus a detailed study of the directivity of the jet noise will be conducted. Acoustic data will be
obtained at several angles with respect to the jet flow (assuming no turning). Flow visualization will be utilized to
determine the final jet turning angle. Most likely tufts will be used, however if the lab PIV system is available and
can be set up in the test facility, some PIV images of the flow field may also be acquired. Jet Mach numbers and
slot heights will be the same as those used in the high-aspect ratio jet study. This data will also be compared with
predicted jet directivity values from jet theory.
4.3 Aerodynamic Verification of Optimized CCW
Once the many possible CCW configurations and blowing conditions have been reduced to a few
promising low noise configurations, aerodynamic performance measurements will be made. Since no acoustic
research has been performed on CCW configurations in the past, it was not known if available aerodynamic data had
been obtained for the best acoustic configurations. Thus, a plan was initiated to re-test the acoustically optimized
CCW configurations and the conventional comparison systems for their aerodynamic performance. This would not
only provide a repeated set of data for conditions tested in previous studies, but would also provide new data for
conditions that had not been tested previously. As noted above, some initial indications from the acoustic tests
indicated smaller slot heights to be acoustically beneficial. This is one case where the aerodynamic tests will
produce new data. In previous studies with these CCW configurations data was only obtained for slot heights down
to 0.006 in. These tests will be done in the Model Test Facility (MTF) at GTRI. They will be done at the same
freestream, jet, and angle of attack conditions as the initial acoustic tests. The models will be mounted on a 6 degree
of freedom balance system, which will give total force and moment values for each condition. Surface pressure
measurements will also be made since the models have pressure taps located along the chord (center span) and will
provide the pressure distribution for the wing. Pressure taps along the span that will also be monitored to verify that
Appendix B
B - 22
2-dimensionality is maintained. These data will provide lift, drag, and moment curves that can be used to determine
conditions of equal lift for different test configurations. This is important for this study because the end result is to
evaluate the acoustic performance of a CCW configuration that could replace a conventional system. Thus it must
provide the same lifting capability, and at most the same amount of drag as the conventional wing. The lift and drag
curves will be used to determine a set of equal lift conditions that will become the desired test points for a more in-
depth acoustic study. For each acoustic test condition, the different configurations will have the same lifting
characteristics and can be compared directly. This will also provide some extra data to fill in the gaps in previous
aerodynamic studies on CCW configurations.
4.4 Initial Results
4.4.1 Acoustic Measurements of Existing CCW Configurations
Much of the initial study of the optimization of the CCW configurations (described in section 4.2.1) has
been completed. Measurements were taken with a B & K 4135, ¼ in. microphone. The microphone was located
11’6” from the trailing edge of the wing, 90 degrees to the freestream flow. Measurements were taken at freestream
velocities of 100, 150, 200, and 250 ft/s (nominal). All configurations were tested at geometric angles of attack of
0, 7, and 14 degrees. The majority of the data presented in this section was acquired at a geometric angle of attack
of 0 degrees and at the highest freestream velocity of about 240 ft/s unless otherwise noted. The trends are the same
at other speeds and angles; so only one case will be shown in order to be brief.
Based on past aerodynamic studies, the best overall aerodynamic characteristics were obtained with the
small CCW flap configurations. The small deflectable flap allowed for low drag during cruise, but by blowing over
the curved upper surface with the flap deflected, significant flow turning could still be achieved when desired. The
highest lift configuration was found to be with the flap deflected 90 degrees. Figure 9 shows acoustic spectra for
several slot velocities with no freestream flow. It shows the basic jet scaling property developed for round jet. For
the measured velocities the theory predicts about a 19 dB increase between the two most extreme cases, which is
similar to that measured (about 16 dB).
This shows that the majority of the noise is associated with the jet noise from the slot and not due to
internal model and facility noise associated with the blowing system above 2 kHz. However, below 2 kHz the
scaling is not followed in the data. This is most likely due to internal noise that is generated from the flow into the
wing on its way to the slot. This essentially contaminates the signal making the noise higher for the lower slot
velocities, but not affecting the higher velocities as much because the jet noise begins to dominate (this is suspected
as the problem, but will be clarified by high aspect ratio nozzle tests). Thus, the difference between the data is less
than predicted by the theory. This is supported by figure 10. This figure shows two slot heights, and hence two slot
areas, at the same slot velocity. However, inside the wing the areas in the flow path remain the same. Since the
mass flow into the wing must be the same as the mass flow out, the doubling of the exit area causes a doubling of
Appendix B
B - 23
the mass flow at the exit, and hence a doubling of the mass flow inside the wing. However, since all the areas inside
the wing are constant, the velocity must double inside the wing in order to double the mass flow. Thus, if noise is
dominated by the internal noise it should follow a sixth power of the velocity scaling as this noise is expected to be
dipole like in nature. If so, the data should reflect a 24 dB increase. However if the noise is dominated by
externally produced jet mixing noise, then it will change only to the extent that the exit area has changed. This will
provide for the jet mixing noise intensity proportional to slot exit area. This translates into a 3 dB increase in noise
after shifting the spectrum for h = 0.006 in. to the left over the spectrum for h = 0.012 in. by a factor of one octave to
allow for the shift in the noise frequencies proportional to a characteristic length. This number is somewhat smaller
than the observed difference in the SPL’s of the two spectra in figure 10. All of these arguments assume that we
can apply the lessons learned from round jets to very high aspect-ratio jets. We reserve our full judgement until
additional studies have been carried out on the HARN nozzle, which will be tested in an acoustically clean facility.
Yet, since the noise increase is not of the order of 24 dB, it can be said that internal noise is not significant in this
case. The fact that the observed difference in spectral SPL’s is more than the expected 3 dB could also be associated
with the scrubbing noise of the CCW slot jet moving over the rounded edge. If so, it is genuinely produced outside
and is not a contribution of internal noise.
Only below about 1500 Hz the data may be contaminated by noise generated internal to the wing. A
muffler was built and installed in the supply line downstream of all valves to eliminate as much upstream noise as
possible. However, due to the small thickness of the wing, inlets into the wing plenum are smaller than desired.
This results in a relatively high velocity flow entering into the plenum with no space to absorb the noise generated.
There are also other internal-noise generators that do not create an aerodynamic problem, but could likely be the
cause of some internal noise generation.
The two major contributors are the paths from the inlet plenum to the slot plenum, and the jack screws that
set the slot height. Two plenums were made to allow the incoming flow to settle and distribute along the entire span
before exiting the slot. Figure 11 shows the flow path through the wing. The flow first comes into the inlet plenum
and then proceeds through rectangular sharp edged slots to the slot plenum. The sharp edges and relatively higher
velocity through the slot can generate noise. Downstream of this, very near the exit of the slot are the jack screws
that allow for slot height adjustment. There are two sets, jack screws that come up from the bottom surface and pull
down screws that thread into the bottom surface. At several spanwise locations, a jack screw and a pull down screw
are positioned in line with the chord line. These look like threaded cylindrical columns to the flow as it passes by on
its way to the slot. These can cause significant shedding, especially since they are located in the contraction area of
the slot where the velocity is accelerated much above the low velocity in the plenum. Similar data is seen at
different freestream velocities and angles of attack.
It is believed that these noise sources may be causing a majority of the noise below 1 kHz and is therefore
the reason the noise is not following the jet prediction method. Budget constraints do not allow us to build an
entirely new model designed for low noise upstream of the slot, hence for the time being this will be noted and data
below 1 kHz will be disregarded as either somewhat corrupted by internal noise or not understood until HARN data
Appendix B
B - 24
becomes available. This will be verified by the HARN data to make sure that it is not associated with a high aspect
ratio jet.
Figure 12 (a -- 0 to 80 kHz, b -- 0 to 20 kHz) shows the noise spectra for several slot jet velocities at a
constant freestream velocity and constant slot height of 0.003 in. There are several things to note. First, with no
blowing there is a large-amplitude well-defined tone. It is also important to note that in general the very low
frequency noise (~ f < 4 kHz) is much greater compared to the data in figure 9. Some of this is from the tunnel
noise itself ( below about 500 Hz) but most of it is flow noise associated with the freestream flow around the wing.
The tone is believed to be due to the shedding of vortices off the bluff trailing edge of the deflected flap. Notice that
blowing, even at low slot jet velocities, significantly reduces the magnitude of the tone. However in this case it is
not completely eliminated and in fact still dominates the spectra at all blowing velocities. The next result to see
from this figure is the jet noise characteristics. At frequencies above about 4 kHz the noise appears to follow the
velocity scaling as described with the data in figure 9. It should also be noted that there is significant noise
associated with the slot flow well past 40 kHz. As expected, the addition of a mean flow did not change the noise
that appears to be noise internal to the wing.
However, the tone mentioned above was unexpected. This presented a problem since the tone dominated
the spectrum at all blowing conditions, thus any gain found in using the CCW over a conventional wing would be
lost if the flap were deflected to 90 degrees.
Since the tone produced with the 90 degree flap was still dominated the noise spectrum even with blowing,
it was decided that reducing the flap deflection might produce a less dominant tone, but still provide enough lift with
the right amount of blowing to equal that of a conventional wing. The flap deflection was reduced to 30 degrees.
Data for similar conditions to the 90 degree deflection are shown in figures 13a and 13b. Again, with no blowing
the tone is present. However, with small amounts of blowing the tone is completely eliminated. Extra tests were
conducted and it was found that as little as 1 psig plenum pressure (Vs ~ 300 ft/s) was enough to completely
eliminate the tone at this flap angle. Since this configuration showed more promise, the remaining parameters were
optimized using the 30 degree flap configuration. Both slot height and slot flow velocity were examined.
The effect of slot height was investigated in these tests. Figure 14 (a and b) shows data with similar
freestream conditions but different slot heights. It is important to note in this case that for the same Cµ, the slot
velocity will be different since Cµ is dependent on mass flow from the slot. Since the goal is to compare the same
lift, it is best to look at the data holding Cµ constant since the same Cµ will give the same lift increment as long as
the flow remains attached in both cases. Thus the data in figure 14 a and b show that there is a lower noise from the
larger slot heights for a given lifting condition. This makes sense since Cµ is proportional to mass flow through the
slot. By increasing the slot height but maintaining the same mass flow (and hence same Cµ) the velocity of the slot
is lower. As was noted in the earlier discussion, decreasing the velocity of a jet flow can significantly decrease the
jet noise associated with it. At this point it appeared that the most appropriate conditions comparing a CCW system
to a conventional system had been found.
Appendix B
B - 25
4.4.2 Determining an “equal lift” condition
The next step was figuring out how to compare the two lift augmentation systems. Aerodynamic data from
previous studies was used in this initial comparison to avoid a long process of swapping models back and forth
between tunnels to narrow down the right acoustic and aerodynamic conditions. Once the model was set up, a large
test matrix could be done without significantly increasing the test time. In view of the fact that setting up the model
is very time consuming it was decided to leave the model set up in the acoustic test tunnel and use previously
obtained aerodynamic data to estimate the lift for the test models. Aerodynamic data was available for both
conventional wings and the CCW. This was provided in the form of lift curves (cl vs α curves) for all the models.
This was convenient since for a CCW, a given Cµ will generally provide a ∆cl over the entire angle of attack range
(not including the extreme high jet velocities and large slots where the jet separates from the surface). Thus, once
the lift for the baseline airfoil was found, this could be compared to the cl for the conventional airfoil and the
needed ∆cl could be calculated by subtracting the two values. This ∆cl was then used to determine the Cµ needed to
match lift provided by the conventional wing flap system. Essentially each Cµ is like a flap setting which shifts the
baseline lift curve by a given amount. For the particular configurations, a Cµ of about 0.04 produced about the same
amount of lift as the conventional wings.
4.4.3 Acoustic Measurements of Conventional Wings
Two conventional wings were tested, one with a 30 degree flap spanning the entire span of the wing (figure
15), and one with a flap deflected 40 degrees spanning the entire wing except for a cut-out region in center span
(figure 16 and photo installed in figure 4). This gap simulates a gap in a flap system on a conventional wing that
may be present to prevent engine exhaust impingement when extended or for some structural reason, as seen on the
Lockheed L-1011 in figure 17. These wings are the same basic airfoil shape as the circulation control wing. The
wings were tested at the same flow conditions as the CCW. Once acoustic data were obtained for these wings old
aerodynamic data were investigated to determine for what blowing conditions the CCW would match the lift
coefficient of the conventional wing. It was found that a Cµ of 0.04 (for CCW at the same angle of attack as the
conventional wing) was sufficient to provide the same lift as the conventional wings.
Initially the conventional wing with the 30 degree flap was tested. Figures 18a and 18b show a comparison
between the conventional wing with the 30 degree flap and the CCW configuration with lowest noise for the
equivalent lift case. In the range between 1 kHz and 10 kHz the CCW has noise levels similar to those of the
conventional system for the CCW h ~ 0.012 in. data. The other two slot heights are also shown, but notice that the
noise levels are higher than the conventional system. Unfortunately neither is a desired result, however it does at
least provide assurance that using the CCW system does not increase the noise to the environment in its minimum
noise configuration.
However, a conventional wing with a flap spanning the entire span of the wing is not really seen in
practice. Most aircraft have a gap in flaps across the span (again see figure 17). This is often due to structural
Appendix B
B - 26
constraints or to prevent engine exhaust from impinging on an extended flap in the case of wing mounted engines.
This difference most likely contributes a fair share of noise to a conventional wing system since flap edge noise is
identified as a major contributor to airframe noise. Thus, this wing was missing a major contribution to the wing
noise problem, a noise source that would most likely be missing, or much reduced in a CCW system. Since the
CCW flap is much smaller, there is no need for a gap in the flap to avoid engine exhaust. Its small size would also
in many cases possibly reduce the need for gaps due to structural concerns. Thus the CCW system with a full span
flap is not unreasonable. In fact, it has even been shown that CCW blowing could be used instead of an aileron for
roll moment, just by varying the blowing between the wings.
For this reason, the single slotted Fowler flap with a gap in the center was installed on the conventional
wing. Acoustic tests were performed on the new configuration similar to the previous tests. Figure 19 shows the
spectra for the two conventional wings at similar freestream conditions. Notice that the two configurations have two
differences, the flap deflection angle and the gap in the flap. Unfortunately a direct comparison of the same flap
angle could not be done because of the different mounting devices for each flap. Note that over a large frequency
range the gap in the flap wing produced much more noise. This could be due to the increased flap deflection angle
or the gap in the flap. This is most likely not due to the change in angle of deflection of the flap. This is based on
the angle of attack results for the conventional wings. It was found that the noise spectrum changed little when the
angle of was changed from 0 to 14 degrees. It is assumed that a similar change in the flap angle (only 10 degrees)
would also have little affect on the noise.
Figures 20a and 20b show the comparison of the gapped wing with the CCW. As expected the gap in the
flap increased the noise on the conventional system significantly and shows a clear advantage to using a CCW
system in the region of 1 kHz to 20 kHz. The CCW system does make more noise at higher frequencies, however
this is not a concern since it is above the hearing limit of humans. Similar results can be seen at other freestream
and angles of attack, however, the magnitude of the difference varies some depending on the conditions. It is worth
noting that the larger slot heights of 0.006 in. and 0.012 in. produced the best results. It is expected that even larger
slot heights would help further. Thus in future experiments some larger slot heights will be tested as well. One
must keep in mind, however, that the configurations have not been rigorously compared in an aerodynamic sense
yet. It is expected that the previous results will be confirmed, but that is the next step.
It should also be noted that this is not the only comparison that can be made. Thus far only comparison of
equal lifting cases with the same angle of attack have been shown. But the CCW system can vary lift with blowing
and angle of attack. It was thought that by comparing the two systems for the same lift with the CCW system at a
different (lower) angle of attack might also produce an advantageous acoustic situation. In this case, if the CCW
were at a lower angle of attack, a higher Cµ would be required. It now must account for the ∆cl of the flap, and the
∆cl due to the ∆α between the two systems. This significantly increased the slot jet noise from the CCW. However,
as was briefly mentioned above the noise from the conventional wing changed little as angle of attack was varied.
Thus, at the angles of attack this was attempted, there was no acoustic advantage to using a reduced angle of attack
CCW to compare with a conventional wing at the same lift condition.
Appendix B
B - 27
Up to this point, only data from a microphone located 11.5 ft from the wing at Θ = 90 degrees has been
shown. This is only part of the noise picture, the changes in directivity of the noise between the two systems must
be compared as well. For this purpose a traverse system was set up in the AFSF to take data at any location in the
fly-over plane between Θ = 30o to Θ = 90o. Data was acquired at 30o, 60o, 90o. Since the traverse translates
linearly, it cannot remain on the same polar arc for each angle, thus it was necessary for comparison purposes to
scale the data using the r squared law for farfield noise propagation. Figure 21 shows the data acquired for the
conventional wings. There appears to be only small changes in the signal with angle. Figure 22 shows similar data
for the CCW system. It should be noted that there are some differences depending on the angle. One must keep in
mind that there may be several factors involved. Note that the 60o and 90o positions do not actually have a line of
sight path to the slot exit which is located on the top surface of the wing. It is also worth noting that the jet from the
slot leaves the trailing edge of the wing at about 56o. Even with freestream velocity, the jet stays relatively close to
that angle for some time beyond the trailing edge of the wing. Figure 23 shows two pictures of tufts taped to the top
surface of the wing. The first picture shows the wing without any blowing, while the second pictures shows a
blowing “on” condition. Notice in figure 23b how the tufts do not bend immediately into the freestream direction
after leaving the trailing edge.
Figures 24, 25 and 26 compare the data for the two wing systems. At 30o the increased noise seen on the
CCW system produces no real advantage over a conventional system. However there is still some noise reduction in
favor of the CCW system at the other two angles. These results indicate that a CCW system certainly has potential
for reducing airframe noise. The results also show some trends in the data thus far, however there is still much left
to study and resolve before all the aspects of the circulation control wing noise issues are solved and helpful to
design of a practical low noise CCW system.
4.5 Contributions to the State of the Art
There are several aspects of this study that will contribute to the general understanding of jet acoustics.
Until now there have been no attempts to directly compare the acoustic noise from a circulation control system and a
conventional wing system. This will provide information for the designer that might want to incorporate circulation
control as a high lift system. In addition to some direct comparisons, the study will provide basic information on
extremely high aspect ratio jets. It will confirm whether or not the same scaling and prediction methods used for
round jets can be applied in this case. This study will also look into the directivity of a jet turned by a surface.
Appendix B
B - 28
5.0 REFERENCES
1 Englar, Robert J., Smith, Marilyn J, Kelley, Sean M. and Rover, Richard C., III. “Development of Circualtion
Control Technology for Application to Advanced subsonic Transport Aircraft,” AIAA paper 93-0644,
presented at AIAA Aerospace Sciences Meeting, January, 1993.
2 Englar, Robert J., Applegate, Constance A. “Circulation control - A Bibliograhpy of DTNSRDC Research
and Selected Outside References: January 1969 through December 1983,” David W. Taylor Naval Ship
Research and Development Center, DTNSRDC-84/052, 1984.
3 Dunham, J. “A theory of circulation Control by Slot-Blowing Applied to a Circular Cylinder,” Journal of
Fluid Mechanics, Vol. 33 Part 3, pp 495-514, 1968.
4Henderson, Campbell. “An Engineering Method for Estimating the Aerodynamic Characteristics of
Circulation Control wings (CCW),” Naval Air Development Center, NADC-82186-60, June 1982.
5Englar, R.J. and Huson, G.G. “Development of Advanced Circulation Control Wing High Lift Airfoils,” AIAA
paper 83-1847, presented at AIAA Applied Aerodynamics Conference, July, 1983.
6 Nichols, J.H. Jr, Englar, R.J., Harris, M. J., Huson, G.G. “Experimental Development of an Advanced
Circulation Control Wing System for Navy STOL Aircraft,” AIAA paper 81-0151, presented at AIAA
Aerospace Sciences Meeting, January, 1981.
7 Englar, Robert J. “Low-Speed Aerodynamic Characteristics of a Small, Fixed Trailing-Edge Circulation
Control Wing Configuration Fitted to a Supercritical Airfoil,” David W. Taylor Naval Ship Research and
Development Center, DTNSRDC/ASED-81/08, March 1981.
8Pugliese, A. J., and Englar, R. J. “Flight Testing the Circulation Control Wing,” AIAA paper 79-1791,
presented at AIAA Aircraft Systems and Technology Meeting, New York, August, 1979.
9Nichols, J.H., Jr. “Development of High Lift Devices for Application to Advanced Navy Aircraft,” Report
DTNSRDC-80/058, AD A084-226, April, 1980.
10 Wilkerson, Joseph B. “An Assessment of circulation Control Airfoil development,” David W. Taylor Naval
Ship Research and Development Center, Report 77-0084, August 1977.
11 Reader, Kenneth R. “Hover Evaluation of the circulation Control High Speed Rotor,” David W. Taylor
Naval Ship Research and Development Center, Report 77-0034, June 1977.
12 Lighthill, M.J. “On sound Generated Aerodynamically. I. General Theory,” Proceedings of the Royal Society
A 211, pp 564-578, 1952.
13 Lighthill, M.J. “On sound Generated Aerodynamically. II. Turbulence as a Source of Sound,” Proceedings of
the Royal Society A 222, pp 1-21, 1954.
Appendix B
B - 29
14 Lighthill, M.J. “Jet Noise,” North Atlantic Treaty Organization Advisory Group for Aeronautical Research and
Development, Report 448, 1963.
15 Lighthill, M.J. “Jet Noise,” AIAA Journal, Vol. 1 Number 7, 1963.
16 Ahuja, K.K. “Correlation and Prediction of Jet Noise,” Journal of Sound and Vibration, 29 (2), pp 155-
168, 1973.
17 Ahuja, K.K., and Bushell, K.W. “An Experimental Study of Subsonic Jet Noise and Comparison with Theory,”
Journal of Sound and Vibration, 30 (3), pp 317-341. 1973.
18 Ahuja, K.K., Tanna, H.K., and Tester, B.J. “An experimental Study of Transmission, Reflection and
Scattering of Sound in A Free Jet Flight Simulation Facility and Comparison with Theory,” Journal of Sound
and Vibration, 75 (1), pp 51-85, 1981.
19 Chaplin, Harvey R., “Wind-Tunnel Investigation of a Small-Scale Two-Dimensional Jet-Flap Wing Model
Over a Large Rand of Jet Deflections,” Navy Department AERO Report 929, TED No. TMB AD-3220.
October, 1957.
20 Salikuddin, M. Brown, W.H., and Ahuja, K.K. “Noise from a Circulation Control Wing with Upper Surface
Blowing,” Journal of Aircraft, Vol. 24 No. 1, January, 1987.
21 Englar, Robert J. “Experimental Investigation of the High Velocity Coanda Wall jet Applied to bluff Trailing
Edge circulation control Airfoils,” David W. Taylor Naval Ship Research and Development Center, Technical
Note AL-308, June 1973.
22 Englar, Robert J. “Low Speed Wind Tunnel Investigation of circulation contol Wing and CCW/Vectored
Thrust STOL Configurations,” Lockheed-Georgia Company, LG86ER0026, march 1986.
23 Englar, Robert J. “Further development of Pneumatic Thrust-Deflecting Powered-Lift Systems,” Journal of
Aircraft, Vol. 25 No. 4, pp324 – 333, April 1988.
24 Lane, Paul Jr. “Ground Controls,” Racecar Engineering, Vol. 9 No. 8, pp 20-23, October 1999.
Appendix B
B - 30
Figure 1: Schematic of CCW concept.
Appendix B
B - 31
Figure 2a: Schematic of CCW flap wing configuration. Base airfoil shape is LG1616
0.5"
1.25" radius
Figure 2b: CCW hinged flap configuration for current CCW model.
Appendix B
B - 32
Figure 3: Schematic of anechoic flight simulation facility
VsΘ
=0
o
Θ =
9
0o
r
Θ =
6
0o
Θ =
3
0o
V
Collect
or
Ope
n J
et
Acoustic
Foam
Appendix B
B - 33
Figure 4: Photo of conventional wing mounted in the anechoic flight simulation facility.
Appendix B
B - 34
0
50
100
150
200
250
300
600 800 1000 1200 1400 1600 1800 2000
Velocity at various motor RPM for the free flight facility,
comparison of "Free flight Mystery formula" and pitot probe inserted at open jet exit.
7-9-99
Free flight formula (Pwall, and Pchamber)
Pitot probe data
Open Jet Velocity, ft/s
.
motor RPM
Figure 5: Comparison of two different methods of velocity calulation for the AFSF.
X
300 psi Supply
Regulating
Valve Muffler
Two Inlets to W ing
3/4" Flexible Tubing
2" High Pressure
Piping 1" High Pressure
Piping
Mass Flow
Venturi
Figure 6: CCW blowing system configuration.
Appendix B
B - 35
Figure 7: Schematic of anechoic static jet facility.
Θ
=
0o
r
=
10
ft
.
Θ
=
30
o
Colle
ctor
Aco
us
ti
c
Fo
am
Θ
=
60
o
Θ
=
90
o
P
l
en
um
Θ
=
1
20o
Appendix B
B - 36
30"
30"
Anechoic
Chamber
Plenum
Round to
Rectangular
Section
Hich Aspect Ratio Nozzle
(Top View)
2.75" h ~ 0.000" - 0.25"
Sliding knife edges
to make adjustable slot
Side View
30"
Flange fabricated to be
compatiable with existing
4" round to 2.75" square
nozzle section
Figure 8a: Schematic of high aspect ratio nozzle (HARN).
Figure 8b: Side view of HARN with CCW flap.
Appendix B
B - 37
30
40
50
60
70
80
0 5 10 15 20
Comparison of various slot velocities. V T = 0.0 ft/s, h ~ 0.006"
(test8-20, T4)
Vs = 660 ft/s
Vs = 875 ft/s
Vs = 1013 ft/s
Vs = 1158 ft/s
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 90o flap
V8 scaling predicts ~19 dB increase
Figure 9: CCW blowing system noise spectra with no freestream flow.
Appendix B
B - 38
20
30
40
50
60
70
0 10 20 30 40 50 60 70 80
Comparison of various slot heights for the same V s = 660 ft/s.
VT = 0.0 ft/s, Θ = 90ο, r = 4.58 ft.
(test11-15, Fp04, gp04)
h ~ 0.006"
h ~ 0.12"
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 10a: Comparison of different slot heights for the same slot velocity.
20
30
40
50
60
70
0 5 10 15 20
Comparison of various slot heights for the same V s = 660 ft/s.
VT = 0.0 ft/s, Θ = 90ο, r = 4.58 ft.
(test11-15, Fp04, gp04)
h ~ 0.006"
h ~ 0.12"
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 10b: Comparison of different slot heights with the same slot velocity.
Appendix B
B - 39
Leading
edge of
wing
Inlet Plenum
Slot Plenum
Inlet
Inlet Rectangular
flow channels
Figure 11: Schematic of CCW blowing flow path inside the current CCW test model.
Appendix B
B - 40
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of various slot velocities. V T = 220.0 ft/s, h ~ 0.006"
90o microphone located at r = 11.5 ft.
(test8-20, T4)
Vs = 0 ft/s
Vs = 660 ft/s
Vs = 860 ft/s
Vs = 1008 ft/s
Vs = 1112 ft/s
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 90o flap
Figure 12a: CCW with 90 degree flap with freestream flow, full spectrum to 80 kHz.
20
40
60
80
100
120
0 5 10 15 20
Comparison of various slot velocities. V T = 220.0 ft/s, h ~ 0.006"
90o microphone located at r = 11.5 ft.
(test8-20, T4)
Vs = 0 ft/s
Vs = 660 ft/s
Vs = 860 ft/s
Vs = 1008 ft/s
Vs = 1112 ft/s
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 90o flap
Figure 12b: CCW with 90 degree flap with freestream spectra from 0 to 20 kHz.
Appendix B
B - 41
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of various slot velocities. V T = 230.0 ft/s, h ~ 0.006"
(test11-15, F)
Vs = 0 ft/s
Vs = 660 ft/s
Vs = 880 ft/s
Vs = 1016 ft/s
Vs = 1116 ft/s
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
V8 scaling predicts ~18 dB increase
Figure 13a: CCW with 30 degree flap with freestream, full spectra from 0 to 80 kHz.
20
40
60
80
100
120
0 5 10 15 20
Comparison of various slot velocities. V T = 230.0 ft/s, h ~ 0.006"
(test11-15, F)
Vs = 0 ft/s
Vs = 660 ft/s
Vs = 880 ft/s
Vs = 1016 ft/s
Vs = 1116 ft/s
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 13b: CCW with 30 degree flap in freestream spectra from 0 to 10 kHz.
Appendix B
B - 42
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of various slot heights for the same C µ = 0.05.
VT = 230.0 ft/s, Θ = 90ο, r = 11.5 ft.
(test11-15, bp20, Fp10, gp06)
h ~ 0.003"
h ~ 0.006"
h ~ 0.012"
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 14a: CCW with 30 degree flap at 3 different slot heights, full spectra from 0 to 80 kHz.
20
40
60
80
100
120
0 5 10 15 20
Comparison of various slot heights for the same C µ = 0.05.
VT = 230.0 ft/s, Θ = 90ο, r = 11.5 ft.
(test11-15, bp20, Fp10, gp06)
h ~ 0.003"
h ~ 0.006"
h ~ 0.012"
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 14b: CCW with 30 degree flap at 3 different slot heights, spectra from 0 to 20 kHz.
Appendix B
B - 43
c = 8.0"
~3
"
30
"
Basic airfoil shape: LG1616 superciritical airfoil.
Figure 15: Schematic of conventional wing with 30 degree full span flap.
2"
c = 8.0 "
3"
30"
Mounting plate
Basic airfoil: LG1616 supercritical airfoil
Figure 16: Conventional wing with 40 degree gapped flap.
Appendix B
B - 44
Figure 17: Example of gap typical in current flap systems on aircraft.
Appendix B
B - 45
20
30
40
50
60
70
80
90
0 10 20 30 40 50 60 70 80
Comparison of CCW to conventional wing wth 30 o flap, Θ = 90o, r = 11.5 ft.
VT = 220.0 ft/s, α = 0o, Cµ ~ 0.04 gives equivalent lift condition
(test8-27, T1, test11-15, bp14, fp10, gp05)
Conventional wing with 30 o flap
CCW, h~ 0.003", Cµ = 0.038
CCW, h~ 0.006", Cµ = 0.045
CCW, h~ 0.012", Cµ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Con. wing, 30 o flap
CCW, 30o flap
Figure 18a: Comparison of CCW and conventional wing with 30 degree flap at similar flight conditions.
20
30
40
50
60
70
80
90
0 5 10 15 20
Comparison of CCW to conventional wing wth 30 o flap, Θ = 90o, r = 11.5 ft.
VT = 220.0 ft/s, α = 0o, Cµ ~ 0.04 gives equivalent lift condition
(test8-27, T1, test11-15, bp14, fp10, gp05)
Conventional wing with 30 o flap
CCW, h~ 0.003", Cµ = 0.038
CCW, h~ 0.006", Cµ = 0.045
CCW, h~ 0.012", Cµ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Con. wing, 30 o flap
CCW, 30o flap
Figure 18b: Comparison of CCW and conventional wing with 30 degree flap at similar flight conditions, from 0 to 20 kHz.
Appendix B
B - 46
30
40
50
60
70
80
90
0 10 20 30 40 50 60 70 80
Comparison of conventional wings. V T = 220.0 ft/s, α = 0o,
(test8-27, T1, test8-31, g)
Conventional wing with 30 o flap
Conventional wing with 40 o flap with gap
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Figure 19: Comparison of the two conventional wings.
Appendix B
B - 47
20
30
40
50
60
70
80
90
0 10 20 30 40 50 60 70 80
Comparison of CCW to conventional wing wth 40 o flap with gap, Θ = 90o,
r = 11.5 ft. VT = 220.0 ft/s, α = 0o, Cµ ~ 0.04 gives equivalent lift condition.
(test8-31, g4, test11-15, bp14, fp10, gp05)
Conventional wing with 40 o flap with gap
CCW, h~ 0.003", Cµ = 0.038
CCW, h~ 0.006", Cµ = 0.045
CCW, h~ 0.012", Cµ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Con. wing, 40 o flap
with gap CCW, 30o flap
Figure 20a: Comparison of CCW and conventional wing with flap gap, 0 to 80 kHz.
30
40
50
60
70
80
90
0 5 10 15 20
Comparison of CCW to conventional wing wth 40 o flap with gap, Θ = 90o,
r = 11.5 ft. VT = 220.0 ft/s, α = 0o, Cµ ~ 0.04 gives equivalent lift condition.
(test8-31, g4, test11-15, bp14, fp10, gp05)
Conventional wing with 40 o flap with gap
CCW, h~ 0.003", Cµ = 0.038
CCW, h~ 0.006", Cµ = 0.045
CCW, h~ 0.012", Cµ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Con. wing, 40 o flap
with gap CCW, 30o flap
Figure 20b: Comparison of CCW and conventional wing with flap gap, 0 to 20 kHz.
Appendix B
B - 48
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of various Θ for the conventional wing. V T = 240.0 ft/s
(test11-17, k, L, m)
Θ = 30o
Θ = 60o
Θ = 90o
SPL - 20Log(1/r)
(Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Con. wing, 40 o
with gap
Figure 21: Comparison of conventional wing noise for different angles with respect to the flow (Θ).
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of various Θ for the CCW. VT = 240.0 ft/s
(test11-17, k, L, m)
Θ = 30o
Θ = 60o
Θ = 90o
SPL - 20Log(1/r)
(Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
CCW, 30o flap
Figure 22: Comparison of CCW noise for different angles with respect to the flow (Θ).
Appendix B
B - 49
(a)
(b)
Figure 23: (a) Tufted CCW in freestream flow without blowing. (b) Tufted CCW in freestream with blowing, slot velocity
~ 600 ft/s, h ~ 0.012 in.
Appendix B
B - 50
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of CCW with conventional wing with 40 o flap with gap. Θ = 30 o
VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift, r = 8 ft.
(test11-14, jp06, 11-17, k)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 30o
Con. wing, 40 o
with gap CCW, 30o flap
Figure 24a: Comparison of CCW and conventional wing at Θ = 30o.
20
40
60
80
100
120
0 5 10 15 20
Comparison of CCW with conventional wing with 40 o flap with gap. Θ = 30 o
VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift, r = 8 ft.
(test11-14, jp06, 11-17, k)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 30o
Con. wing, 40 o
with gap CCW, 30o flap
Figure 24b: Comparison of CCW and conventional wing at Θ = 30o, 0 to 20 kHz.
Appendix B
B - 51
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of CCW with conventional wing with 40 o flap with gap. Θ = 60 o,
r = 5 ft. VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift
(test11-14, hp06, 11-17, m)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.039
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 60o
Con. wing, 40 o
with gap CCW, 30o flap
Figure 25a: Comparison of CCW and conventional wing at Θ = 60o.
20
40
60
80
100
120
0 5 10 15 20
Comparison of CCW with conventional wing with 40 o flap with gap. Θ = 60 o,
r = 5 ft. VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift
(test11-14, hp06, 11-17, m)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.039
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 60o
Con. wing, 40 o
with gap CCW, 30o flap
Figure 25b: Comparison of CCW and conventional wing at Θ = 60o, 0 to 20 kHz.
Appendix B
B - 52
20
40
60
80
100
120
0 10 20 30 40 50 60 70 80
Comparison of CCW with conventional wing with 40 o flap with gap, Θ = 90o,
r = 4.63 ft. VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift
(test11-14, hp06, 11-17, m)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 90o
CCW, 30o flap
Con. wing, 40 o
with gap
Figure 26a: Comparison of CCW and conventional wing at Θ = 90o.
20
40
60
80
100
120
0 5 10 15 20
Comparison of CCW with conventional wing with 40 o flap with gap, Θ = 90o,
r = 4.63 ft. VT = 240.0 ft/s, Cµ = 0.04 produces equivelant lift
(test11-14, hp06, 11-17, m)
Conventional wing with gap
CCW, 30o flap, h~0.012", C µ = 0.040
SPL (Pref = 20 X 10-6 Pa)
Frequency, kHz ( ∆f = 32 Hz)
Θ = 90o
CCW, 30o flap
Con. wing, 40 o
with gap
Figure 26b: Comparison of CCW and conventional wing at Θ = 90o, 0 to 20 kHz.
Appendix B
B - 53
Appendix B
B - 54
