• Tiada Hasil Ditemukan

THE EXPERIMENTAL STUDY OF THE EFFECT OF BIAS FLOW ON AERODYNAMIC FORCES OF NACELLE

N/A
N/A
Protected

Academic year: 2022

Share "THE EXPERIMENTAL STUDY OF THE EFFECT OF BIAS FLOW ON AERODYNAMIC FORCES OF NACELLE "

Copied!
50
0
0

Tekspenuh

(1)

i

THE EXPERIMENTAL STUDY OF THE EFFECT OF BIAS FLOW ON AERODYNAMIC FORCES OF NACELLE

D-CHAMBER

By

MOHAMMAD AMIR IZUDDIN ISHAK (Matrix no: 120386)

Supervisor:

Dr. Mohd Azmi Ismail

JUNE 2017

This dissertation is submitted to Universiti Sains Malaysia

As partial fulfilment of the requirement to graduate with honors degrees in BACHELOR OF ENGINEERING (MECHANICAL ENGINEERING)

School of Mechanical Engineering Engineering Campus Universiti Sains Malaysia

(2)

i

ACKNOWLEDGEMENTS

Firstly, I would like to thank my supervisor, Dr. Mohd Azmi Ismail for his guidance throughout the whole project. I would like to thank him a lot for his advice and ideas that he gave to me during my fabrication and experimental process. In addition to that, I would also like to express my gratitude to Qummare Azam for his encouragement and help from him to understand the project which enables me to carry out the analysis more efficiently.

Besides, I would also like to express my gratitude to all the lecturers and technical staffs in the School of Mechanical Engineering and School of Aerospace Engineering, Universiti Sains Malaysia, especially Dr. Norizham Razak, Mr Baharom Awang, Mr Ridwan Yusuf and Mr Mahmud Isa for sharing their though and experience with me. This has certainly helped me in completing the project more effectively.

Last but not the least, I would like to extend my gratitude to the persons that always give moral support to me throughout the whole project. The persons that I mentioned are my father, my mother, my siblings and all my friends in School of Mechanical Engineering.

Thanks a lot for your moral support and without all of you, this project might have not finished by the given time.

(3)

ii

Title of thesis: THE EXPERIMENTAL STUDY OF THE EFFECT OF BIAS FLOW ON AERODYNAMIC FORCES OF NACELLE D-CHAMBER

Date of submission (Academic year): 7th June 2017 (2016 / 2017) Candidate (Matrix no.): 120386

Name of supervisor: Dr. Mohd Azmi Ismail

DECLARATION

This work has not previously been accepted in substance for any degree and is not being concurrently submitted in candidature for any degree.

Signed ... (Candidate) Date ...

Statement 1

This thesis is the result of my own investigation, except where otherwise stated. Other sources are acknowledged by giving explicit references. Bibliographies /references are appended.

Signed ... (Candidate) Date ...

Statement 2

I hereby give consent for my thesis, if accepted, to be available for photocopying and for interlibrary loan, and for the title and summary to be made available outside organizations.

Signed ... (Candidate) Date ...

(4)

iii

Table of Contents

ACKNOWLEDGEMENTS ... i

DECLARATION ... ii

LIST OF FIGURES AND GRAPH ... v

LIST OF TABLES ... vi

NOMENCLATURE ... vi

ABSTRAK ... vii

ABSTRACT ... viii

INTRODUCTION ... 1

1.0 Introduction ... 1

1.1 Problem Statement ... 4

1.2 Aim and Objectives ... 5

1.3 Scope of Research ... 5

LITERATURE REVIEW ... 6

2.1 Liner’s Sound Abatement ... 6

2.3 Liner’s Heat Transfer ... 9

2.4 Lift & Drag ... 13

METHODOLOGY ... 15

3.1 Model Development ... 15

3.2 Fabrication Process ... 17

3.3 Test Setup and Measurement ... 19

3.4 Experimental Procedure ... 22

(5)

iv

RESULT & DISCUSSION ... 25

4.1: Modified Bias Model ... 25

4.2 Comparative of D-chamber model ... 30

CONCLUSION ... 33

6.1 Conclusion ... 33

6.2 Future Work ... 34

REFERENCES ... 35

APPENDICES ... 38

Sample Calculation ... 38

Data Compilation ... 39

(6)

v

LIST OF FIGURES AND GRAPH Page

Fig 1.1 Piccolo System 2

Fig 1.2 Acoustic Liner 2

Fig 1.3 Double degree bias flow liner 3

Fig 1.4 Location of bias liner on nacelle D-chamber 4

Fig 2.1 Progress in aircraft noise reduction 6

Fig 2.2 Helmholtz resonator 7

Fig 2.3 Grazing-bias flow interaction 8

Fig 2.4 Translating of liner for varying impedance 9

Fig 2.5 Schematic flow pattern through a hole 10

Fig 2.6 Dimension of test sample 11

Fig 2.7 Numerical solution of honeycomb back porosity by Ives 12 Fig 2.8 Numerical comparison of ice protection system 12

Fig 2.9 Micro Blowing Technique (MBT) arrangement 14

Fig 2.10 Numerical solution of pressure contour of nacelle D-chamber 14

Fig 3.1 Falcon 20G 16

Fig 3.2 Nacelle D-chamber model 16

Fig 3.3 Hydraulic Press Brake machine 18

Fig 3.4 D-chamber models’ frame 18

Fig 3.5 Perforated hole dimension on back facing sheet 18

Fig 3.6 Open loop wind tunnel 19

Fig 3.7 D-chamber attachment on 3D Balancing Unit 20

Fig 3.8 1.5 kW reciprocating compressor 21

Fig 3.9 Pnematic linkage from compressor to D-chamber 21

Fig 3.10 Digital Manometer 22

Fig 3.11 Lift & drag force data display 22

Fig 3.12 Experiment setup of nacelle D-chamber 24

Fig 3.13 Schematic drawing of experiment setup 24

Fig 4.1 Lift & drag force against pressure difference at 10 m/s 25

Fig 4.2 𝐶𝐿 and 𝐶𝐷 at 10 m/s 28

(7)

vi

Fig 4.3 𝐶𝐿 and 𝐶𝐷 28

Fig 4.4 Lift & drag force of both models 29

Fig 4.5 𝐶𝐷 against free stream at 𝛼 = 6° 30

Fig 4.6 𝐶𝐿 against free stream at 𝛼 = 6° 31

LIST OF TABLES Page

Table 3.1 Properties of the developed models 15

Table 3.2 Aluminium 1100 properties at 25 °C 17

Table 4.1 Variation of 𝐶𝐷 & 𝐶𝐿 at certain parameters 27

Table 4.2 Percentage difference of 𝐶𝐷 & 𝐶𝐿 31

NOMENCLATURE

A area (𝑚2)

𝛼 angle of attack (°)

𝑉 free stream velocity (𝑚/𝑠) 𝜌𝑎𝑖𝑟 air density (𝑘𝑔 𝑚⁄ 3) 𝐹𝐷 drag force (N) 𝐹𝐿 lift force (N)

∆P pressure difference (Pa) 𝐶𝐷 coefficient of drag 𝐶𝐿 coefficient of lift 𝐹𝐵 blowing fraction

(8)

vii

ABSTRAK

Sistem perlindungan ais dan liner bias adalah satu konsep bersepadu yang boleh digunakan untuk menyediakan pesawat kedua-dua pengurangan bunyi dan sistem perlindungan ais.

Kebanyakkan pesawat semasa menggunakan sistem Tiub Piccolo Anti-Ais yang menghentam jet udara panas untuk perlindungan ais. Liner bias adalah pelapik akustik yang membolehkan udara mengalir melalui setiap sel-sel liang dan bertindak sebagai peranti pengurangan pelbagai bunyi. Jika udara panas dibuang ke luar melalui sel-sel ini, ia boleh berfungsi sebagai aliran berat sebelah. Kesan aliran berat sebelah ke arah prestasi aerodinamik nasel ruang-D dikaji.

Siasatan aliran berat sebelah telah dijalankan dengan menggunakan dua model maju nasel ruang-D; satu model pepejal dan satu model diubahsuai. Model diubahsuai mempunyai kulit yang berliang dengan satu barisan tunggal lubang berliang pada bahagian bawah.

Bekalan aliran berat sebelah adalah diukur dengan perbezaan tekanan antara tekanan dalaman ruang-D dengan tekanan ambien sekitarnya, sehingga 160 Pa antara 10-20 m/s halaju aliran bebas.

Menurut keputusan eksperimen, aliran berat sebelah benar menunjukkan keputusan yang menyakinkan. Aliran ini membantu dalam mengurangkan pekali seretan dan meningkatkan pekali angkat pada nasel ruang-D. Ujikaji juga menemui bahawa tekanan kritikal pengurangan drag adalah 60 Pa dan peratusan perbezaan bagi pekali seret dan pekali angkat tidak dipengaruhi oleh hadlaju aliran bebas.

(9)

viii

ABSTRACT

The ice protection system and bias liner is an integrated concept that could be used to provide both aircraft ice protection and noise abatement. Current aircraft uses a Piccolo Tube Anti-Icing system which impinges jets of hot air for ice protection. The bias liner is an acoustic liner which allows air to flow through each of the cells and act as sound abatement device. If the hot air is dumped overboard via these cells, it can function as a bias flow. The effect of bias flow towards the aerodynamic performance of nacelle D- chamber is studied.

An investigation was carried out by using two developed models of nacelle D-chamber;

one solid model and one modified bias model. The modified model is having a perforated skin with single row of porous holes at the bottom part. The bias flow supply was positively quantified by the internal pressure difference of the nacelle D-chamber with the ambient, up to 160 Pa ranging from 10-20 m/s free stream velocities.

According to the experimental study, bias flow does provide a promising result. The flow helps in reducing drag coefficient and enhance the lift coefficient of the nacelle D-chamber.

The experiment also found that critical pressure of drag reduction is 60 Pa and the percentage of difference for lift coefficient and drag coefficient are independent with free stream velocity.

(10)

1

CHAPTER ONE INTRODUCTION

1.0 Introduction

An aircraft usually accumulates ice when flying through clouds at higher altitudes.

Ice is formed from super-cooled water droplets in the clouds impacting on the aircraft surfaces where the surface temperature is 0°C and below. Ice accretion on the aircraft wings can cause the aircraft to divert or make the aircraft difficult to maneuver and the worst part is it can cause the aircraft to crash.[1] This is due to the ice accretion changing the shape of the aircraft hence result in a reduction or the loss of aerodynamic performance. Also, the loss is due to the formation of non-aerodynamic shapes of the aircraft, precisely on leading wing and engine nacelle. This ice accretion also can contribute on higher fuel consumption as the higher drag force is generated when these shapes are present.[2] Therefore, protection against icing is required.

An ice protection system is usually installed on these critical surfaces. Modern commercial aircraft uses hot air which is bled from the aircraft’s engine compressor and then impinged it on the back of those critical surfaces that requiring ice protection. The hot air impingement method is commonly named as Piccolo Tube Anti-Icing (PTAI). This system is one of the most popular hot air ice protection system utilized in leading wing and nacelle lipskin due to its higher efficiency and reliability[3]. Figure 1.1 shows the piccolo system as one of the ice protection system of an aircraft.

Due to the current and projected concerns about community noise pollution from commercial air traffic, huge number of innovative noise reduction concept solutions are being considered and one of them is acoustic liner. Typically, there are two types of liner that can be installed inside the engine nacelle which are acoustic liner and bias acoustic liner. Acoustic liner act as sound reducer while bias acoustic liner can gives more advance properties of sound abatement. However, acoustic liner treatments for engine nacelles have been an effective means of suppressing turbomachinery noise for last decades. Figure 1.2

(11)

2

shows a conventional acoustic liner 20 consisting of a honeycomb structure 14 bonded using adhesive 10 between a perforated sheet 16 and a non-perforated sheet 12.[4]

Fig 1.1: Piccolo System[5]

Fig 1.2: Acoustic Liner

(12)

3

Conventional acoustic liner cannot be installed on every part of the nacelle and one significant area is the nacelle lip or commonly named as nacelle D-chamber. This is due to the ice protection system will destroys the liner caused by hotspot issues.[6] In the other hand, bias liner which utilizes bias flow as a means of changing liner impedances for better noise abatement have a better heat transfer mechanism compared to conventional acoustic liner.[7] The bias liner is almost similar with the conventional acoustic liner with the exception that the liner’s backing sheet is also perforated, which allows air to flow through the liner. For the bias liner to perform as a part of ice protection integration system, it should be performing substantial enough in term of heat transfer mechanism as the piccolo system.

A bias flow is the introduction of airflow, blowing or suction, perpendicular to the liner as seen in figure 1.3.

According to Federal Aviation Administration (FAA), aircraft noise is a serious cause of environmental noise pollution and the only way to deal with it is to mitigate these noises. Therefore, the FAA has developed a program named, Continuous Lower Energy, Emissions and Noise (CLEEN), which FAA's principal environmental effort to accelerate the development of new aircraft and engine technologies and advance sustainable alternative jet fuels. CLEEN projects develop technologies that will reduce noise, emissions, and fuel burn and enable the aviation industry to expedite integration of these technologies into current and future aircraft. CLEEN is a venture between Boeing, General Electric, Honeywell, Pratt & Whitney and Rolls-Royce which contributed in this fund- sharing program for the past 6 years.[8]

Fig 1.3: Double degree bias flow liner [9]

(13)

4

1.1 Problem Statement

It is a big step for aircraft manufacturer company to consider the integration of an ice protection system with sound abatement device inside the nacelle D-chamber. The integration would be a better way to ensure there is no ice accretion happens on the nacelle lip, also reducing the sound production from the act of the impingement of hot air to be dumped overboard[10]. The hotspot and non-uniform skin temperature distribution issues produced by the piccolo system also can be resolved by the integration system.

Anthony Ives from Queens’ University of Belfast had proven that the bias acoustic liner has a better heat transfer rate compared to the conventional acoustic liner. The bias acoustic liner also shown a promising result of heat transfer coefficient similar as the piccolo system. For the hotspot issue, the bias liner allows the trapped hot air in the cell to heat up small areas of the D-chamber which produces a more uniform distribution of temperature on the nacelle skin.[11] However, he only considered that the bleed hot air will be dumped overboard via exhaust vent as seen in figure 1.1.

Thus, this study is concerning about a concept of having a perforated surface on the D-chamber skin, just underneath the bias liner location as shown in figure 1.4. The reason is to let the hot bleed air exhausting through this perforated surface instead of normal exhaust vent and act as bias flow. However, there is uncertainty about the aerodynamic performance of the nacelle D-chamber on perforated surface with bias flow, which is the subject of this thesis.

Fig 1.4: Location of bias liner on nacelle D-chamber

(14)

5

1.2 Aim and Objectives

The primary goal of this study is to develop an understanding of the aerodynamic performance of nacelle D-chamber with the effect of bias flow. Following are the objective of this study:

• Develop two models of nacelle D-chamber; solid model and modified bias model.

• Generate the lift and drag coefficient from experimental data

• Comparison of aerodynamic performance between models

1.3 Scope of Research

In the real scenario, the hot air is impinging into back of nacelle D-chamber. The hot air is generated from the aircraft’s engine compressor and bled out through the exhaust vent. This hot air impingement will heat-up the nacelle skin thus protect the nacelle from any ice accretion. The air that been dumped overboard via perforated skin also might be affecting the aerodynamic mechanism of the nacelle. However, this project is only focusing on the understanding of lift and drag behavior. Thus, the implementation of hot air is negligible.

The thesis contains five chapters. Chapter 2 gives a review of the research work on the bias liner as well as related topics which could give an understanding of the aerodynamic mechanisms involved. Chapter 3 gives a brief experimental assessment of nacelle D-chamber aerodynamic performance and technical aspect of test rig. Chapter 4 describes the experimental results of the nacelle D-chamber models. Chapter 5 concluded the main findings of the experiment results and reviews what further research can be done for the nacelle D-chamber bias flow model.

(15)

6

CHAPTER 2

LITERATURE REVIEW

The most significant augmentation of research in bias liner has concentrated on the noise abatement properties of the liner and the heat transfer mechanism. The early section of this literature review looks at the fundamental of the bias liner’s noise abatement properties.

The second section reviews the research of heat transfer mechanism of the bias liner. The third section looks at the blowing effect would have on the aerodynamics performance. The final section looks at recent studies onto the drag and lift of the nacelle D-chamber.

2.1 Liner’s Sound Abatement

Engine noise is one of the major contributors to the overall sound levels as aircraft operate near airports due to the usage of turbofan. In which these engines are commonly used on commercial transports due to their advantage for higher performance and lower noise.[12] Therefore, NASA has conducted and sponsored research aimed with noise reduction for commercial aircraft as seen in figure 2.1, especially as the growth of the air traffic and the impact of noise on the community will increase if the noise levels are not reduced.

Fig 2.1: Progress in aircraft noise reduction

(16)

7

Number of ways for sound abatement had been proposed. One the way is using acoustic liner. This liner usually installed in cowl zones and exhaust nozzles of jet engines as the liner’s purpose is to provide absorption of unwanted acoustic energy. The modern acoustic liner is based on the principle of the Helmholtz resonator. The Helmholtz resonator can be said like a chamber of trapped air that connected to a long and small duct as seen in figure 2.2.

When the air inside the neck of the resonator is resonating, two effects can be produced; either the sound is amplified or dissipated. Thus, for this acoustic liner case, it will produce slower sound when the air produced from the engine passing through the liner and act as a sound abatement system. The liner is made of perforated face sheet, solid back sheet, and honey comb sandwich core with septum to abate the noise.[13]

Bias acoustic liner is invention based on Dean’s study about blowing air passed through a porous plate, which increasing the acoustic resistance. The only different with an acoustic liner and a bias liner is that bias liner has a perforated back sheet, compared to solid back sheet found in acoustic liner. This porous back sheet allows air to pass through the liner.[14]

The work of Sun et al shows that grazing flow across the front plate also have effect on acoustic resistance. It is shown that there is some difference between grazing-in flow and grazing-out flow interactions as figure 2.3 shows outflow and inflow arrangement for grazing-bias flow interaction.

Fig 2.2: Helmholtz resonator

(17)

8

For outflow, acoustic resistance drops below the non-grazing flow at high bias flow speed whereas the inflow has the effect of increasing the acoustic resistance in porous plate or orifice and acoustic resistance of an orifice increases slowly with the increase of bias flow speed. However, acoustic resistance is dominated by the bias flow effect when bias flow speed is generally high speed. [15]

Fig 2.3: Grazing-bias flow interaction a) Outflow

b) Inflow

Mathematical study of acoustic waves by Lighthill shows the bias flow present and absent demonstrates some changes: (i) the increase of the order of the differential equation from two to three (ii) the acoustic pressure as a function of the distance from the wall (iii) bias flow can change significantly the acoustic pressure in the boundary layer. The almost same finding was obtained by the study of Eldredge et al where effectiveness of the bias liner absorbing axial acoustic waves in a circular duct was found to be strongly influenced by (i) duct circumferential and cross-sectional area (ii) midpoint length (iii) open area ratio of facing porous plate. [16],[17]

(18)

9

Ahuja and Gaeta have also studied on details of varying of perforate geometry and porosity by translating one perforate over another equally in porosity as an effort to vary the acoustic liner’s impedance as seen in figure 2.4 They also found that such as star-shaped orifices, resulted in higher resistance and thus higher absorption at lower incident sound pressure levels but did not show a significant difference in impedance at higher incident sound pressure levels. [18]

Fig 2.4: Translating of liner for varying impedance

2.3 Liner’s Heat Transfer

The major research of heat transfer of porous liner was primary for the usage in cooling of combustor wall and turbine blades. Sparrow had experimentally investigated the heat transfer for porous plates positioned at stiff angles to an oncoming flow. In his findings, the present flow field has some features in common with a stagnation flow of a 13 and 19 tube holes and the use of the pitch size on the characteristic dimension in the Nusselt number is quite effective in eliminating the dependence of the results on the pitch- to-diameter ratio. More recently, Dorignac’s study shown that his results were compared quite well with the experimental case by Sparrow on large perforation diameters. However, as the pitch-to-diameter ratio increases, this correlation started to deviate with the experimental result. [19],[20]

(19)

10

Cho experimentally investigated the local heat transfer coefficient for both the windward and leeward regions near the hole by using a naphthalene sublimation technique.

At the windward surface, the local transfer coefficient increases rapidly as the flow approaches the hole entrance due to acceleration of the flow resulting in a thinning boundary layer. For the leeward surface, the mass transfer is very low due to a low entrainment velocity on the surface compared to the exit jet velocity. He also found that the peak heat transfer occurred within the hole at the point of reattachment, region B in figure 2.5. Overall, peak heat transfer inside the hole is about four times the heat transfer given by fully developed tube flow. [21]

Fig 2.5: Schematic flow pattern through a hole

(20)

11

Ives experimentally and numerically investigated the bias liner’s heat transfer mechanism. His test sample of liner was constructed from aluminum facing and backing sheets of 1 mm thickness and an aluminum honeycomb core of approximately 22 mm thickness. He also used honeycomb cells have a width of approximately of 6.35 mm and a wall thickness of 0.3 mm. For the backing sheet, a hole pattern of 0.7 mm hole diameter at a pitch of approximately 11.2 mm along the breath and approximately 6.5 mm hole pitch along the length is drilled and that’s give a backing sheet porosity of 1% for one hole per honeycomb cell. His test sample dimension and numerical solution are shown in figure 2.6 and 2.7 respectively.

In his finding, numerical and experimental results shown that the bias liner has a higher heat transfer coefficient compared to conventional acoustic liner. Also, Ives’s numerical result gives a lower heat transfer coefficients for the bias liner than that of the piccolo system for both multi holes and single hole as seen in figure 2.8. However the heat transfer coefficients are still acceptable to provide ice protection [11],[22]

Fig 2.6: Dimension of test sample (in mm)

(21)

12 Fig 2.7: Numerical solution of honeycomb of single hole back porosity

Fig 2.8: Numerical comparison of ice protection system

(22)

13

2.4 Lift & Drag

One of the major aircraft performance parameters analyzer are drag and lift capability. It also can be used to evaluate other major aircraft characteristics such as range, climb rate and maximum speed. Lift force is defined as the component of the total aerodynamic force normal to the flow direction. While drag force is defined as the component of the total aerodynamic force parallel to the flow direction.

There are three different ways of test that can be used to determine the lift and drag force: (i) wind tunnel test (ii) empirical method and analytical correlations and (iii) Computational Fluid Dynamics. However, the test number (ii) is not suitable because it cannot define the actual design of nacelle lofting and flow at null angle of attack.[23]

The location of nacelle on the wings also contributed to the aircraft’s performance.

This was studied by Landrum which she found that any changes of the orientation of nacelle on the wing has only small effect on the drag and lift coefficient of the whole aircraft.

However, for an aircraft with two nacelles, the effect of changing the orientation of nacelle will give larger effect on the drag and lift coefficient. [24]

Back in 1996, the influence of porous holes on aerodynamic performance was studied by Mineck and Hartwich. Their findings included test on full cord porosity of NACA 0012 airfoil. In general, porosity did have a different outcome. One of the outcome was reduction on lift-curve slope and addition on drag when the airfoil was tested at different Mach Number[25].

However, different result exhibited when the porous holes are having small jets due to air pass through. This method is called Micro Blowing Technique (MBT) and was made famous by Danny P. Hwang. MBT is proven to reduce turbulent skin friction up to 50 % in subsonic flow. Figure 2.9 shows the typical arrangement of the MBT. The outer layer can be represented as the bias liner in the anti-icing and sound abatement integration system.[26]

(23)

14

Azam numerically investigated the drag performance of the nacelle lip. He compared the drag coefficient between nacelle lip with and without bias flow. According to the simulation results, bias flow helps in reducing drag coefficient of nacelle lip as meant of boundary layer thickness variation as shown in figure 2.10. The highest drag coefficient percentage difference happens at free stream velocity of 5 m/s. However, the percentage of drag coefficient decreased as the free stream velocity increased.[27]

Fig 2.9: Micro Blowing Technique (MBT) arrangement

Fig 2.10: Numerical solution of pressure contour of nacelle D-chamber

(24)

15

CHAPTER 3 METHODOLOGY

This section describes the model development, experimental setup and testing used for the experiment. Two models of nacelle D-chamber are developed with one the model idealised as the solid model and the other one is for the bias flow model. The bias flow model had perforated surface for air exhausting system. The models are then tested and measurements of drag and lift forces are taken to calculate drag and lift coefficient.

3.1 Model Development

Nacelle D-chamber is annular shape that conform the most front face of an aircraft engine nacelle intake. In this experimental, the model is developed based on Dassault Falcon 20G aircraft nacelle design. This business jet is built by Dassault Aviation, a French aviation company. Figure 3.1 shows the actual figure of nacelle used for the aircraft and the actual figure of the aircraft.

The air pressure inside the chamber is assumed to be very similar along the circular chamber and the circular shape of the double curvature nacelle D-chamber were replaced with a simplified straight section of nacelle D-chamber with a single curvature as shown in figure 3.2. The experimental results of this simplified geometry were expected to have a little deviation from that of the actual design. Some previous research works have concurred with this assumption.[27], [28]

For the experiment, two models were developed; 1 solid model and 1 modified bias model. Both models have the same geometric shape that based on the aircrafts’ nacelle lip.

Table 3.2 shows the properties of the developed models.

Solid Modified

Dimension (mm) 115 x 270 115 x 270

Mass (g) 259 238

No. of holes Frontal Area (𝒎𝟐)

NA 0.03105

110 0.03105 Table 3.1: Properties of the developed models

(25)

16

The forces data collected from the experiment is going to be interpreted as a set of dimensionless numbers. These numbers are coefficient of drag, 𝐶𝐷 and coefficient of lift, 𝐶𝐿. The formulae are given

𝐶

𝐷/𝐿

= 𝐹

𝐷/𝐿

1

2

𝜌

𝐴𝑖𝑟

. 𝑉

2

. 𝐴

(1) Where

𝐹

𝐷/𝐿 is drag or lift force

𝜌

𝐴𝑖𝑟 is density of air 𝑉 is free stream velocity

𝐴

is models’ frontal area

Fig 3.1: Falcon 20G

Fig 3.2: Nacelle D-chamber model

(26)

17

3.2 Fabrication Process

Two models are developed for the experiment which one solid model and one modified bias model with perforated skin for bias flow implementation. Aluminium 1100- Alloy sheet of thickness 1 mm has been chosen for the skin due to its excellent forming characteristics. The properties of aluminium alloy is shown in table 3.2 The bulk sheet was cut into two small size of 420 x 270 mm sheets by using Hydraulic Press Brake machine.

This machine is available in both School of Aerospace Engineering and School of Mechanical Engineering. For the shape profile model, acrylic perspex of 30 mm thickness are chosen. This acrylic perspex act like a main frame of the model. Each model consists of two equal size of acrylic perspex that been hold by 345 mm long aluminium shaft. This 10mm diameter of shaft act like a bone that hold the two acrylic perspex in place, also used for connecting part for attaching the model with a balancing unit that will be explained in experiment setup. Figure 3.3 shows the cutting activity of aluminium done by an assistant engineer. These aluminium shafts are fabricated using lathe machine.

Next step was to wrap the aluminium sheet around the models’ frame. This is done by pre-wrap the aluminum sheet and ensure the sheet fits the frame. An adhesive glue is used for sticking the aluminum sheet together with the frame, also to avoid any air leakages due to sticking process. Figure 3.4 shows the actual models’ frame and comparison with original made by previous project. For the perforated bias flow models, the same processes applied with addition of a hole drilled for pneumatic pipe fitting at the side of the model.

However, numbers of 110 holes with 1 mm diameter were drilled at the lower region of the modified model. These holes dimension and pitch size are based on the experimental study done by Ives for bias acoustic liner as shown in figure 3.5.

Density (x1000 kg/𝒎𝟐) 2.71 Elastic Modulus (GPa) 70-80 Tensile Strengh (MPa) 110

Poisson’s Ratio 0.33

Elongation (%) 12

Thermal Conductivity (W/m-K) 218 Table 3.2: Aluminium 1100 properties at 25 °C

(27)

18 Fig 3.3: Hydraulic Press Brake machine

Fig 3.4: D-Chamber models’ frame

Fig 3.5: Perforated hole dimension on back facing sheet (in mm)[11]

(28)

19

3.3 Test Setup and Measurement

The next step is experimental test setup. For this experiment, we are using an open loop subsonic wind tunnel. This wind tunnel can provide free stream velocities up to 30 m/s with test section size of 300mm x 300mm x 600mm as seen in figure 3.6. The D- chamber model will be placed inside the test section by attachment with a balancing unit.

This 3D Electronic Balancing unit will measure the lift, drag and pitch moment of the D- chamber.

(i)

(ii)

Fig 3.6 Open loop wind tunnel (i) Overall view

(ii) Test section

(29)

20

Both D-chambers each will be attached to the balancing unit and placed inside the test section. The balancing unit uses 3 different load cells as sensors to measure the force.

For precaution, this balancing unit will be calibrated once per month by respective personnel. Also, this balancing unit allow us to change the angle of attack (α) if necessary however we are only focusing for drag measurement at zero angle as in real scenario where the orientation of the nacelle is on zero angle. This setup is as seen in figure 3.7. For data record, the balancing unit is connected to a computer for data acquisition and representation. A software of AFA50 TQ Wind Tunnel Data Acquisition is utilized for the force measurement and wind tunnel speed measurement and display as seen in figure 3.11.

A 1.5 kW compressor is needed for supplying the bias flow of air as seen in figure 3.8. These reciprocating-piston type of compressor supplies the bias flow into D-chamber and exhausting through the porous hole region on the bottom of the D-chamber. This flow is supplied from the side of the D-chamber model. A complete set for pneumatic setup of M8-standard size linked from the compressor to the D-chamber is shown in figure 3.9. The pressure inside the D-chamber is controlled by the opening and closing of compressor’s valve. The measurement is taken by a digital manometer as pressure sensor. This manometer is capable to measure pressure up to 78.85 kPa or equivalent to 11 Psi. The manometer has two pressure transducer inlets as shown in figure 3.10. However, only one inlet is used for the experiment and placed on the D-chamber.

Fig 3.7: D-chamber attachment on 3D Balancing Unit

(30)

21

Fig 3.8: 1.5 kW reciprocating compressor

Fig 3.9: Pnematic linkage from compressor to D-chamber

(31)

22

Fig 3.10: Digital manometer

Fig 3.11: Lift & drag force data display

(32)

23

3.4 Experimental Procedure

The free stream velocities are varied from 10 m/s to 20 m/s in the wind tunnel. This is done by controlling the rotation speed of the fan motor adjusted to desired free stream velocity. Corresponding to the velocities, the reading of the manometer shown negative values of pressure indicating that the pressure outside the D-chamber is higher than the pressure inside the D-chamber. These negative pressures are considered as ‘ambient’

condition of the free stream. For supplying the bias flow, the compressor’s valve is opened slowly to steadily increase the inside pressure of D-chamber. The valve is opened until the digital manometer shows a zero pressure. At this moment, the internal pressure is equal to the ambient pressure.

The study is divided into two set of experiments. First, to study the aerodynamic performance at various amount of internal pressure. The measurement of the drag and lift forces are recorded for each ascending internal pressure of 10 Pa increment, up to 160 Pa.

This is mean that for every stream velocity, a set of 12 different internal pressure is used for lift and drag forces measurement. However, steps mentioned are only applicable for the perforated nacelle D-chamber model.

Second, to compare the performance of the models with respect to constant nacelle D-chamber’s internal pressure. The pressure will be constant-positively varied for every free stream velocity and force data recorded will be discussed as comparative between the solid model and the modified bias model. The original model only need to through a series of force measurement with different free stream velocities. Each set of parameters of experiment is repeated three times to avoid any measurement errors and uncertainties. The shown data in next chapter are based on average data from the experiment. The exact display of the experiment setup is shown in figure 3.12. A schematic drawing of the experiment setup is shown in figure 3.13.

(33)

24

Fig 3.12: Experiment setup of nacelle D-chamber

Fig 3.13: Schematic drawing of experiment setup

(34)

25

CHAPTER FOUR RESULT & DISCUSSION

This chapter discusses all the results of the experiment. The first section discusses the results of the experiment for lift and drag coefficient of modified bias model at zero angle of attack (𝛼 = 0°). The final section will look at the comparative results of lift and drag coefficient at 6° angle of attack for both models. (𝛼 = 6°). All the tabulated results are presented in Appendix.

4.1: Modified Bias Model

In this part of experiment, only modified bias flow model is used. The model is placed in zero angle of attack (𝛼 = 0°) and the internal pressure is positively varied from ambient pressure throughout numbers of free stream velocity. Figure 4.1 shown the graph of lift and drag forces of the modified bias model at free stream velocity of 10 m/s. The graph later translated into coefficient of drag, 𝐶𝐷 and coefficient of lift, 𝐶𝐿 with equation 1.

Fig 4.1: Lift & drag force against pressure difference at 10 m/s

0 0.1 0.2 0.3 0.4 0.5 0.6

0 10 20 30 40 50 60 70 80 90 100 120 140 160

Force (N)

∆P (Pa)

Lift

Drag

(35)

26

Figure 4.2 shows the variation of 𝐶𝐷 and 𝐶𝐿 of the modified model against internal pressure difference, ∆P at 10 m/s of free stream velocity and at 𝛼 = 0. It is recorded that the external pressure provided by the free stream is -250 Pa. The negative symbol indicates the higher pressure outside the D-chamber. The compressor’s valve is opened until the digital manometer shown a zero value, which mean both external pressure and internal pressure are equalized. (∆P=0)

It is found that for 𝐶𝐷 for 10 m/s free stream, the value drops gradually from initial equalized pressure up to ∆P=60 Pa. The value drops 16.67% from 0.2898 to 0.2415. This result shown that the bias flow on nacelle D-chamber can be assumed as an application of micro-blowing technique.

The technique is meant to reduce the local skin drag as the result of reduction of the gradient of tangential velocity at the surface wall. This tangential velocity on the first-layer grid around the nacelle D-chamber can be represented as the magnitude of local skin- friction. The same solution was found and briefly explained by Gao et al which they numerically study the effect of micro-blowing technique on a super critical airfoil.[29] An almost same explanation regarding the pressure field was given by Azam as described in chapter 2.

However, it is noticed that the 𝐶𝐷 is increased as the ∆P elevated onwards to 160 Pa. This is caused by the pressure-drag penalty produced by the micro-blowing at greater pressure or higher blowing fraction. The blowing fraction, 𝐹𝐵 is a ratio of blowing/bias condition to the glazing condition. Although there is an experimental result by Hwang and Biesiadny showed that with the increase of the pressure outlet or the blowing fraction, the pressure-drag would increase dramatically which also leads to total drag intensification as the total drag is the sum of skin-friction drag and pressure drag.

(36)

27

The mechanism behind the intensify of the pressure drag was not provided by the authors but they did brief some explanation about the total drag intensification after a certain condition. The cause of the total drag intensify was the overwhelming pressure drag elevation compared to skin-friction reduction. At low pressure, the total drag reduction was the result of the skin-friction drag reduction and some local pressure drag reduction. As the pressure rises, the rate of pressure drag elevation starts to defy the rate of skin-friction reduction thus result in total drag intensification.[30]

The same pattern was observed for the free stream velocity of 14 m/s and 15 m/s as in figure 4.3. For 14 m/s, the 𝐶𝐷 dropped 15.9% from 0.3195 to 0.2685 and for 15 m/s, it is dropped 10.9% from 0.3268 to 0.2910. From the mentioned result, the reduction in 𝐶𝐷 is decreasing as the free stream velocity increases. This can be said that as the free stream velocity is getting faster, the effect of bias flow towards total drag reduction is obsolete.

For 𝐶𝐿, there is a steady growth observed as the ∆P increases. This is because the bias flow would also be affecting the lift of the nacelle D-chamber as meant of changing the pressure distributions. The almost same study by the mentioned journal concludes that for a flow field without a shock wave, when the micro blowing zone happened at the lower wall of the airfoil, it can enhance the lift by utilizing the pressure peak and pressure trough.[31] This is the exact same condition applied to the nacelle D-chamber modified bias model. At the maximum supplied pressure difference (∆P=160 Pa), it is recorded the highest value of 𝐶𝐿 for all three free stream velocities.

Both 𝐶𝐷 and 𝐶𝐿 have shown some increment in values as the free stream velocity increases. Free stream 10 m/s recorded 𝐶𝐷 of 0.2898 at the initial moment of experiment.

The values were slightly increased to 0.3195 and 0.3268 at 14 m/s and 15 m/s respectively.

For 𝐶𝐿, the values are 0.0644, 0.0698 and 0.0715 for initial pressure condition at 10 m/s, 14 m/s and 15 m/s respectively.

(37)

28

At the same condition of internal pressure, the 𝐶𝐷 and 𝐶𝐿 are increasing gradually as the free stream velocity increases. The same upward trends are recorded at ∆P=60 Pa and 160 Pa for both 𝐶𝐷 and 𝐶𝐿 at the three different stream velocities. Table 4.1 show the tabulated result of 𝐶𝐷 and 𝐶𝐿 at equalized pressure (∆P =0), critical drag reduction pressure (∆P = 60 Pa) and at maximum supplied pressure (∆P =160Pa). It is noticed that the 𝐶𝐷 had regain its initial value as the ∆P increases from 60 Pa to 120 Pa. This region is where the pressure drag penalty took place in the experiment.

Table 4.1: Variation of 𝐶𝐷 & 𝐶𝐿 at certain condition

Fig 4.2: 𝐶𝐿 and 𝐶𝐷 at 10 m/s

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35

0 10 20 30 40 50 60 70 80 90 100 120 140 160

CL/CD

∆P (Pa)

10 m/s

CL

CD

∆P = 0 ∆P = 60 Pa ∆P = 160 Pa

Free Stream (m/s) 𝐶𝐷 𝐶𝐿 𝐶𝐷 𝐶𝐿 𝐶𝐷 𝐶𝐿

10 0.2898 0.0644 0.2415 0.0858 0.2952 0.1060 14 0.3195 0.0698 0.2685 0.0939 0.3202 0.1074 15 0.3268 0.0715 0.2910 0.0954 0.3291 0.1097

(38)

29 (i)

(ii)

Fig 4.3: 𝐶𝐿 and 𝐶𝐷 (i) 14 m/s (ii) 15 m/s

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35

0 10 20 30 40 50 60 70 80 90 100 120 140 160

∆P (Pa)

14 m/s

CL

CD

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35

0 10 20 30 40 50 60 70 80 90 100 120 140 160

CL/CD

∆P (Pa)

15 m/s

CL

CD

(39)

30

4.2 Comparative of D-chamber model

For this part of experiment, both models are placed in 6° angle of attack (𝛼 = 6°). For modified bias model, the internal pressure is maintained at 30 Pa higher than ambient throughout the various stream velocities. The percentage of reduction and addition of 𝐶𝐷 and 𝐶𝐿 are recorded respectively. Figure 4.3 shown the variation values of lift and drag force across free stream velocities.

Fig 4.4: Lift & drag force of both models

Both models behave proportionally with the change of stream velocity. As the velocity gone faster, the forces are increasing linearly with different manner. For solid model, it holds the highest trendline of drag force followed by lowest trendline of lift force, represented by brown line and blue line respectively. Bias model in contrast, shown a reduction in drag force trendline and hold a slightly higher lift force trendline compared to solid model. Both line represented by yellow line and grey line respectively.

0 1 2 3 4 5 6 7 8 9

10 11 12 13 14 15 16 17 18 19 20

Force (N)

Velocity (m/s)

Lift(Solid) Drag(Solid) Lift(Bias) Drag(Bias)

(40)

31

Figure 4.4 shown a graph of 𝐶𝐷 against stream velocities for both solid model and modified bias model. From the graph, it is seen that the reduction of 𝐶𝐷 is steadily maintain for the modified bias model across the stream velocities. The range for percentage of 𝐶𝐷 reduction is from 50.84-54.97%. The lowest percentage of reduction is occurred at 11 m/s where’s the highest reduction is at 20 m/s. It can be said that the percentage of 𝐶𝐷 reduction is only depending on the internal pressure, not to the free stream velocities. Table 4.2(a) shows the tabulated result for 𝐶𝐷 percentage reduction for the experiment.

The almost same behavior recorded for the 𝐶𝐿 as in figure 4.5. It also can be said that the percentage of improvement in 𝐶𝐿 are independently with the change of stream velocity considering the internal pressure difference is maintained. For the range of percent, the values varied from 20-29.03% as shown in table 4.2(b). As for the mentioned journal, the numerical results of the critical airfoil indicated that a 12.8–16.8% reduction of total drag and 14.7–17.8% increase of lift would be achieved with a blowing fraction, 𝐹𝐵 of 0.05 as its took place at 𝛼 = 5° [29]

Fig 4.5: 𝐶𝐷 against free stream at 𝛼 = 6°

0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1

10 11 12 13 14 15 16 17 18 19 20

CD (Solid/Bias)

Velocity (m/s)

CD (Solid)

CD (Bias)

(41)

32 Fig 4.6: 𝐶𝐿 against free stream at 𝛼 = 6°

0.1 0.11 0.12 0.13 0.14 0.15 0.16 0.17

10 11 12 13 14 15 16 17 18 19 20

CL(Solid/Bias)

Velocity (m/s)

CL (Solid)

CL (Bias)

𝑽 (m/s) Improvement (%)

10 28.57

11 26.92

12 29.03

13 27.89

14 21.91

15 24.00

16 20.69

17 21.21

18 23.29

19 20

20 24.24

𝑽(m/s) Reduction (%)

10 51.28

11 50.84

12 51.05

13 53.87

14 52.32

15 53.02

16 53.83

17 54.01

18 52.79

19 53.41

20 54.97

(a) (b)

Table 4.2: Percentage difference (a) Improvement of 𝐶𝐿

(b) Reduction of 𝐶𝐷

(42)

33

CHAPTER FIVE CONCLUSION

Recalling the aim of the research study was to understand the aerodynamic performance of the nacelle D-chamber with effect of bias flow. The study was based on the integration concept between an ice protection system and a sound abatement device on the nacelle lip. In this chapter, the main findings of the study, the conclusions drawn and the suggestions for potential future work, are summarized.

6.1 Conclusion

The study was started with development of nacelle D-chamber models. Two models of nacelle D-chambers were developed. First model, the solid model was used as the baseline study for understanding the aerodynamic performance in the term of lift coefficient, 𝐶𝐿 and drag coefficient, 𝐶𝐷. Second model, modified bias flow model was used for implementing the bias flow. This model was the key-model to further studied the effect of bias flow. There are two major findings that were found in the study.

First, the bias flow helps in reducing the 𝐶𝐷. However, there was a backlash of the drag reduction. This backlash was identified as the pressure-drag penalty. This penalty is associated with the internal pressure of nacelle D-chamber rise above the critical pressure, which is 60Pa. The 𝐶𝐷 started to intensify and regained its initial value back. Other than that, bias flow also enhanced the lift. As the pressure difference is getting higher, 𝐶𝐿 also is getting better.

Second, it was found that the percentage of either reduction of 𝐶𝐷 or improvement of 𝐶𝐿 does not affected by the stream velocity, as long as the internal nacelle D-chamber pressure is maintained at a certain value. In this case, the pressure is kept 30Pa higher than the ambient throughout the free stream velocities. It is indicated that a 50.58-54.97%

reduction in 𝐶𝐷 and a 20-29.03% improvement in 𝐶𝐿.

(43)

34

In the nutshell, bias flow does give promising results toward the aerodynamic performance of the nacelle D-chamber with a backlash at certain extend. Instead of providing ice protection and sound abatement system, the integration system also can enhance the aerodynamic performance.

6.2 Future Work

As mentioned earlier, the demonstrated experiment was neglecting the present of the hot air as the bias flow. Thus, implementing the hot air in the further study can be more useful to understand the aerodynamic performance of the nacelle as in the real application of ice protection and sound abatement. Also, numerical study to validate the experimental result could be proven the effectiveness of the bias flow. Future study could be also directed at studying the different location of the porous holes that affecting the drag reduction. By this mean, an optimum location for porous hole would be decided both numerically and experimentally. Besides that, different arrangement of porous holes also can be considered as a parameter affecting the effectiveness of the drag reduction.

It is crucial to understand the mechanism of pressure drag intensification that found in the experiment. By doing this experimentally and numerically, it can provide a good resource for further research in the field of drag reduction, especially micro-blowing technique. Other than that, the study of any backlashes that associated with lift also can be conducted regarding the matter.

Finally, the integration system is not only limited to Piccolo Tube Anti-Icing system. The different type of bled hot air for ice protection can also be considered for the concept such as Swirl Anti-Icing. A direct comparison could also be made for those two types of integration system for ice protection and sound abatement.

(44)

35

REFERENCES

[1] F. T. Lynch and A. Khodadoust, “Effects of ice accretions on aircraft aerodynamics,”

Prog. Aerosp. Sci., vol. 37, no. 8, pp. 669–767, 2001.

[2] Y. Cao, Z. Wu, Y. Su, and Z. Xu, “Aircraft flight characteristics in icing conditions,”

Prog. Aerosp. Sci., vol. 74, pp. 62–80, Apr. 2015.

[3] M. A. Ismail and M. Z. Abdullah, “Applying Computational Fluid Dynamic to Predict the Thermal Performance of the Nacelle Anti-Icing System in Real Flight Scenarios,” Indian J. Sci. Technol., vol. 8, no. 30, Nov. 2015.

[4] Moe, J.M., Wunsch, J.J. & Sperling, M.S. 2005, "Method and Apparatus for Noise Abatement and Ice Protection of an Aircraft Engine Nacelle Inlet Lip", US Patent number 20050006529.

[5] Riley, S.J. & James, E.H. 1994, "Influence of various exhaust geometrical parameters on the effectiveness of an aero-engine intake thermal anti-icing system", Proceedings of the International Gas Turbine and Aeroengine Congress and Exposition, Jun 13-16 1994, Publ by ASME, New York, NY, USA, Hague, Neth, pp. 1.

[6] R. Elangovan, R. F. Olsen, and N. D. Reynolds, “Modeling of Acoustically Treated Nacelle Lip Transpiration Flow Anti-Icing System,” 26th Int. Congr. Aeronaut. Sci., pp. 1–13, 2008.

[7] E. Moers, D. Tonon, A. Hirschberg, E. Moers, D. Tonon, and A. Hirschberg, “Sound absorption by perforated walls with bias / grazing flow : experimental study of the influence of perforation angle Sound absorption by perforated walls with bias / grazing flow : experimental study of the influence of perforation angle,” no. April, pp. 3449–3453, 2012.

[8] General Electric, “Enhanced Noise Abatement Climb Final Report,” submitted for Continuous Lower Energy, Emissions and Noise (CLEEN) Program, Federal Aviation Administration (FAA) 2014. DOT/FAA/AEE/2014-05

[9] R. H. Thomas, M. M. Choudhari, and R. D. Joslin, “Flow and noise control: review and assessment of future directions,” Russell J. Bertrand Russell Arch., vol.

NASA/TM-20, no. April, 2002.

(45)

36

[10] S. Raghunathan et al., “Key aerodynamic technologies for aircraft engine nacelles,”

Aeronaut. J., vol. 110, no. 1107, pp. 265–288, May 2006.

[11] A.O. Ives "Perforated Honeycomb Acoustic Liner Heat Transfer" PhD Thesis, Queen's University Belfast, 2009

[12] D. L. Huff, “Technologies for Turbofan Noise Reduction,” 10 th AIAACEAS Aeroacoustics Conf. 1012, no. September, pp. 1--27, 2004.

[13] B. Houston, J. Wang, Q. Qin, and P. Rubini, “Experimental and numerical investigation of Helmholtz resonators and perforated liners as attenuation devices in industrial gas turbine combustors,” Fuel, vol. 151, pp. 31–39, 2015.

[14] Dean, P.D. & Tester, B.J., "Duct Wall Impedance Control as an Advanced Concept for Acoustic Suppression", NASA Technical Report NASA CR-134998, 1975

[15] Sun, X., Jing, X., Zhang, H. & Shi, Y., "Effect of grazing-bias flow interaction on acoustic impedance of perforated plates", Journal of Sound and Vibration, vol. 254, no. 3, pp. 557-573, 2002.

[16] M. J. Lighthill, “On the energy scattered from the interaction of turbulence with sound or shock waves,” Math. Proc. Cambridge Philos. Soc., vol. 49, no. 3, p. 531, Jul. 1953.

[17] J. D. Eldredge and A. P. Dowling, “The absorption of axial acoustic waves by a perforated liner with bias flow,” J. Fluid Mech., vol. 485, no. 6, p. 307-335, 2003.

[18] K. K. Ahuja and R. J. Gaeta, “Active Control of Liner Impedance by Varying Perforate Orifice Geometry,” Nasa Tech. Memo., December, 2000.

[19] E. M. Sparrow and M. C. Ortiz, “Heat transfer coefficients for the upstream face of a perforated plate positioned normal to an oncoming flow,” Int. J. Heat Mass Transf., vol. 25, no. 1, pp. 127–135, Jan. 1982.

[20] E. Dorignac, J. J. Vullierme, M. Broussely, C. Foulon, and M. Mokkadem,

“Experimental heat transfer on the windward surface of a perforated flat plate,” Int.

J. Therm. Sci., vol. 44, no. 9, pp. 885–893, 2005.

[21] H. Cho, M. Jabbari, and R. Goldstein, “Experimental mass (heat) transfer in and near a circular hole in a flat plate,” J. Heat Mass Transf., vol. 40, no. 10, pp. 2431–2443, 1997.

(46)

37

[22] A. Ives, J. Wang, S. Raghunathan, E. Benard, and P. Sloan, “Three Dimensional Numerical Solution of Heat Transfer in a Honeycomb Cell,” in 7th AIAA ATIO Conf, 2nd CEIAT Int’l Conf on Innov and Integr in Aero Sciences,17th LTA Systems Tech Conf; followed by 2nd TEOS Forum, 2007, no. September, pp. 18–20.

[23] L. G. Trapp and H. G. Argentieri, “Evaluation of nacelle drag using Computational Fluid Dynamics,” J. Aerosp. Technol. Manag., vol. 2, no. 2, pp. 145–154, 2010.

[24] Landrum, E.J. "Effect of nacelle orientation on the aerodynamic characteristics of an arrow wing-body configuration at Mach number 2.03", 1996.

[25] R. E. Mineck and P. M. Hartwich, “Effect of Full-Chord Porosity on Aerodynamic Characteristics of the NACA 0012 Airfoil,” April, 1996.

[26] D. Hwang, “Review of research into the concept of the microblowing technique for turbulent skin friction reduction,” Prog. Aerosp. Sci., vol. 40, no. 8, pp. 559–575, 2004.

[27] Q. Azam, M. A. Ismail, N. M. Mazlan, and M. Bashir, “Numerical Comparison of Drag Coefficient between Nacelle Lip-Skin with and without Bias Acoustic Liner,”

Int. Rev. Mech. Eng., vol. 10, no. 6, p. 390, Sep. 2016.

[28] M. A. Ismail, S. H. Mohammad Firdaus, S. N. Soid, M. K. Khalil, and H. Yusoff,

“Numerical Investigation on Uniformity of Heat Distribution of Swirl Anti-Icing System,” Indian J. Sci. Technol., vol. 8, no. 30, Nov. 2015.

[29] Z. Gao, J. Cai, J. Li, C. Jiang, and C.-H. Lee, “Numerical Study on Mechanism of Drag Reduction by Microblowing Technique on Supercritical Airfoil,” J. Aerosp.

Eng., vol. 30, no. 3, p. 4016084, May 2017.

[30] D. P. Hwang, T. J. Biesiadny, “Experimental evaluation of the penalty associated with micro-blowing for reducing skin friction.” Proc., 36th Aerospace Science Meeting and Exhibit, American Institute of Aeronautics and Astronautics, Reston, VA. 1998

[31] C. Junxuan and G. Z. Xun, “Numerical Study on Drag Reduction by Micro- Blowing/Suction Compounding Flow Control on Supercritical Airfoil,” Procedia Eng., vol. 99, no. December 2015, pp. 613–617, 2015.

(47)

38

APPENDICES Sample Calculation

Given:

𝑉 = 10 (𝑚/𝑠) 𝐹𝐷 = 1.95 (N) 𝜌𝑎𝑖𝑟 = 1.2 (𝑘𝑔 𝑚⁄ 3) 𝐴 = 0.03105 (𝑚2) From eq 1:

𝐶𝐷 = 𝐹𝐷 {1

2 𝜌𝐴𝑖𝑟. 𝑉2. 𝐴}

⁄ 𝐶𝐷 = 1.95

{1

2 (1.2). (102). (0.03105)}

⁄ 𝐶𝐷 = 1.0317

(48)

39

Data Compilation

Modified Bias Model 𝑉 = 10 𝑚/𝑠

𝛼 = 0°

𝑉 = 15 𝑚/𝑠 𝛼 = 0°

Pressure Diff Bias Model Coefficient

Lift Drag CL CD

0 0.12 0.54 0.0644 0.2898

10 0.12 0.53 0.0644 0.2844

20 0.14 0.52 0.0751 0.2791

30 0.15 0.5 0.0805 0.2683

40 0.15 0.48 0.0805 0.2576

50 0.16 0.47 0.0858 0.2522

60 0.16 0.45 0.0858 0.2415

70 0.17 0.5 0.0912 0.2683

80 0.18 0.52 0.0966 0.2791

90 0.17 0.53 0.0912 0.2844

100 0.19 0.53 0.1019 0.2844

120 0.18 0.53 0.0966 0.2844

140 0.2 0.55 0.1073 0.2952

160 0.22 0.55 0.1180 0.2952

Pressure Diff Bias Model Coefficient

Lift Drag CL CD

0 0.3 1.37 0.0715 0.3268

10 0.3 1.35 0.0715 0.3220

20 0.32 1.33 0.0763 0.3172

30 0.35 1.34 0.0834 0.3196

40 0.36 1.31 0.0858 0.3125

50 0.37 1.28 0.0882 0.3053

60 0.4 1.22 0.0954 0.2910

70 0.4 1.25 0.0954 0.2981

80 0.4 1.3 0.0954 0.3101

90 0.41 1.32 0.0978 0.3148

100 0.43 1.35 0.1025 0.3220

120 0.44 1.36 0.1049 0.3244

140 0.43 1.38 0.1025 0.3291

160 0.46 1.38 0.1097 0.3291

Rujukan

DOKUMEN BERKAITAN

5.3 Experimental Phage Therapy 5.3.1 Experimental Phage Therapy on Cell Culture Model In order to determine the efficacy of the isolated bacteriophage, C34, against infected

In view of the above phenomenon and to fill-in the gap, this study attempts: first, to determine consumers’ general purchasing behaviour pattern when they

In this research, the researchers will examine the relationship between the fluctuation of housing price in the United States and the macroeconomic variables, which are

Therefore, at 1400C with coarser grain size, the composite mechanical properties slightly decreases but the readings were quite high compared to the composites sintered lower

Hence, this study was designed to investigate the methods employed by pre-school teachers to prepare and present their lesson to promote the acquisition of vocabulary meaning..

Taraxsteryl acetate and hexyl laurate were found in the stem bark, while, pinocembrin, pinostrobin, a-amyrin acetate, and P-amyrin acetate were isolated from the root extract..

The construction of numbers will be started with natural numbers, and then extended to the integers, rational numbers and finally the real numbers...

With this commitment, ABM as their training centre is responsible to deliver a very unique training program to cater for construction industries needs using six regional