khagendra thesis

本科毕业设计论文

题目翼型气动性能实验与共轴旋翼试验台设计

专业名称航空工程

学生姓名 Khagendra Kumar Yadav

指导教师赵旭

毕业时间 2013年 6月

Experimental research on aerodynamic performance of an airfoil and test-rig design of

coaxial rotating wings

Supervisor: Xu Zhao

Thesis submitted to the Northwestern Polytechnical University for the degree of

Bachelor

in Aeronautical Engineering

Candidate: Khagendra Kumar Yadav

June, 2013

Xi’an, china


I

The Assignment of Graduation Thesis

1. Topic

Experimental research on aerodynamic performance of an airfoil and test-rig design of coaxial

rotating wings

2. Guiding ideology and purpose

Through the work of this thesis, the student can be made to consolidate and improve the basic

theories and professional knowledge, and he is trained to use his previous knowledge

comprehensively. The student should master the influence factors and research methods of

aerodynamic performance of coaxial rotor airfoil and be trained and cultivated to work and do

scientific research independently, in order to improve his professional quality.

3. The main technical indicators

(1) Understand the basic knowledge of coaxial rotating wings.

(2) Carry out airfoil aerodynamics performance measurement using wind tunnel and wind

velocity.

(3) Compare CFD results with experimental results to analyze , , carying with AOA

and wind velocity.

(4) Complete concept design of coaxial rotating wings test-rig using mechanical and

aeronautical knowledge.

4. Schedule

(1) Week 3-4, complete wind tunnel test.

(2) Week 5-6, get familiar with the parameters and properties of counter rotating wing airfoil

(3) Week 7-9, get familiar with Fluent, establish the calculation model of the symmetrical

airfoil, master the calculation methods and skills and analyze the results.

(4) Week 9-11, compare CFD results with experimental results.

(5) Week 12-14, complete test-rig design.

(6) Week 14-15, complete the thesis

(7) Week 16, prepare the defense.


II

5. The main reference books and reference materials

[1] Bohorquez, F. DRUM: Rotor Hover Performance and System Design of an Efficient

Coaxial Rotary Wing Micro Air Vehicle . PhD Thesis, University of Maryland, USA,2007.

[2]Jian Tang, Dragos Viieru and Wei Shyy. Effects of Reynolds Number and Flapping

Kinematics on Hovering Aerodynamics . 45th AIAA Aerospace Sciences Meeting and Exhibit

8,Reno, Nevada,2007.

[3]J.C. Bell; M. Brazinskas; S. D. Prior; C. Barlow; M. A. Erbil; M. Karamanoglu. Development

of a Test-Rig for Exploring Optimal Conditions of Small Unmanned Aerial Vehicle Co-

Axial Rotor Systems. Department of Product Design and Engineering, School of Engineering

and Information Sciences, Middlesex University, Trent Park Campus, Bramley Road, London,

N14 4YZ, UK.

[4] John D . Anderson, Jr. Fundamentals of Aerodynamics, McGraw-Hill Education,2005.

[5] Gambit 6.2 User Guide, Fluent Inc. of American,2006.

[6] Fluent 6.2 User Guide, Fluent Inc. of American,2006.

学生: 指导教师: 系主任:


III

摘要

前后对称上凸下凹翼型可以用于旋翼无人机的研发。为了掌握翼型的气动特性，本

文通过低速风洞实验，完成了某翼型的升力系数、阻力系数和压力系数的测量，风速范围

20~40m/s, 攻角为-2°~16°。并采用 CFD 软件 Fluent 模拟翼型在试验工况下的流动特性，

湍流模型采用 realizable k-e 模型。结果表明，数值计算和实验结果吻合较好，说明此翼型

具有较高的升阻比和升力系数。本文还给出简单的误差分析。最后本文对共轴旋翼试验台

进行了初步的概念设计，对比文献中的几种共轴旋翼试验台和优缺点，结合西北工业大学

现有的旋翼试验台，进行了改进和方案设计，为研究无人机悬停状态旋翼气动特性奠定基

础。

关键词：翼型，共轴双旋翼, 无人机, 风洞实验, 旋翼实验台


IV

Abstract Fre-and-aft symmetrical upper surface protrude lower surface concave (USPLSC) airfoils have been developed for the coaxial rotating wing of an UAV. In order to understand the

aerodynamical performance of the airfoil, low speed wind tunnel experiment is carried out to measure lift coefficient, drag coefficient and pressure distribution under wind speed 20~40 m/s. Results are compared with CFD prediction using Fluent software. A satisfied agreement has been

achieved and error analysis is provided. Finally, the concept design of test-rig for coaxial rotating wing is completed based on the comparison of different rigs in literature and the existing rig in

NPU. The purpose of this study is to conduct a parametric investigation on the performance of USPLSC airfoils for coaxial rotating wing UAV. The main objective of this study is to test

USPLSC in wind tunnel and compare the experiment result with CFD simulation result. Through this study, the test-rig is designed for measuring the hovering performance of the coaxial rotating

wing.

Keywords: Airfoil, coaxial rotor, UAV, wind tunnel experiment, rotating wing, test-rig


V

Contents The Assignment of Graduation Thesis .................................................................................................. I

Abstract………………………………………………………………………………………………………………………………………………… IV Nomenclature……………………………………………………………………………………………………………………………………… VII Acronyms………………………………………………………………………………………………………………………………………………VII

Chapter 1 Introduction........................................................................................................................1

1.1 General discussion of UAVs .................................................................................................... 1

1.2 UAV model ............................................................................................................................ 1

1.3 Brief introduction of coaxial counter rotating wing .................................................................. 3

1.3.1 Background and advantages of coaxial rotor.........................................................................5

Chapter 2 Aerodynamics of airfoils ......................................................................................................7

2.1 Airfoil.................................................................................................................................... 7

2.2 Aerodynamic forces ............................................................................................................... 7

2.2.1 Lift and drag .......................................................................................................................8

2.2.2 Velocity and pressure distributions ......................................................................................9

2.3 USPLSC airfoil ...................................................................................................................... 10

Chapter 3 Airfoil pressure measurement............................................................................................ 12

3.1 Experiment equipment ........................................................................................................ 12

3.1.1 Wind tunnel...................................................................................................................... 12

3.1.2 Airfoil model……………………………………………………………………………………………………………….………...13

3.1.3 Pitot rake.......................................................................................................................... 14

3.1.4 Airfoil pitch control ........................................................................................................... 14

3.1.5 Pressure measuring system ............................................................................................... 15

3.1.6 Test section ...................................................................................................................... 15

3.2 Experiment status................................................................................................................ 15

3.3 Calculation methods ............................................................................................................ 16

3.4 Experimental result and analysis........................................................................................... 19

3.5 Error analysis....................................................................................................................... 20

Chapter 4 CFD simulations ................................................................................................................ 22

4.1 Gambit and Fluent ............................................................................................................... 23

4.2 CFD simulation of USPLSC airfoil ........................................................................................... 25

4.2.1 Establishment of CFD model .............................................................................................. 27


VI

4.2.2 Working conditions ........................................................................................................... 28

4.2.3 Comparison of Cl, Cd .......................................................................................................... 28

4.2.4 Comparison of Cp.............................................................................................................. 29

Chapter 5 Test-rig design for coaxial rotor.......................................................................................... 34

5.1 Test-rig Development .......................................................................................................... 36

5.2 Single rotor test-rig at NWPU ............................................................................................... 39

5-3 Development of coaxial rotor test-rig ................................................................................... 40

5.4 Safety devices ..................................................................................................................... 43

Chapter 6 Conclusion and future work ............................................................................................... 44

6.1 Conclusion .......................................................................................................................... 44

6.2 Future work ........................................................................................................................ 44

References ....................................................................................................................................... 46

Acknowledgment.............................................................................................................................. 48

Graduation project summary............................................................................................................. 49


VII

Nomenclature

A = body area (m2) cd = drag coefficient cl = lifting coefficient Cp = pressure coefficient D' = drag force (N) ds = a differential segment of the airfoil surface L/D = Lift to drag ratio L' = lifting force (N) P = static pressure at the point of interest

P∞ = free stream static pressure pT = total or stagnation upstream pressure q = dynamics pressure v = flow velocity (m/s)

v∞ = free stream velocity ρ = density of fluid (kg/m3) (1.225 kg/m3) (dry air)

α = angle of attack

Acronyms

AOA = angle of attack CCRW = coaxial counter rotating wing CFD = computer fluid dynamics GPS = global positioning system LTWT = low turbulence wind tunnel MAV = micro air vehicle NWPU = Northwestern Polytechnical University PLA = people liberation army RPM = revolutions per minute SUAV = small unmanned air vehicles UAV = unmanned air vehicle USPLSC = upper surface protrude lower surface concave

VTOL = vertical takeoff and landing 2D = two dimensional


1

Chapter 1 Introduction

The USPLSC Project is a 2 years research project at NWPU, China, which is focused on in coaxial rotor wing UAVs to obtain vertical takeoff and landing. Fixed wing UAVs have the disadvantage of requiring runway or launcher for takeoff- landing and not being able to hover. On

the other hand, rotary wing UAVs have the advantage of being able to hover, takeoff and land vertically with agile maneuvering capability at the expense of high mechanical complexity.

There are many studies on rotorcraft UAVs with different rotor configurations. In this thesis attention is paid to configure the wing by one USPLSC airfoil, working likes a rotating blade. The design of the fixed wing UAV conversion into the rotor wing UAV modifications

remained the basis of the original aerodynamic design. Chapter 2 presents a brief survey of airfoils. Chapter 3 presents wind tunnel equipment and experiment. Chapter 4 shows the CFD

simulation and result compare with experiment. Chapter 5 presents details and development of the coaxial rotating wing test-rig.

1.1 General discussion of UAVs From the mid-1960s the use of UAVs for intelligence, surveillance and reconnaissance

(ISR) mission has featured in operations over Chechnya, China, the Middle East, South-East Asia and the former Yugoslavia. Currently, more than three dozen nations are active in developing UAV technology, and the leader in advancements of UAV technology is the US.

Over five dozen different programs including the American Predator, Global Hawk and Shadow make up the United States’ arsenal of UAV [1].

In recent years, interest has grown in using UAVs predominantly for military applications, but also used in a small but growing number of civil applications, such as policing, firefighting, and nonmilitary security work, such as surveillance of pipelines. UAVs are often preferred for

missions that are too "dull, dirty, or dangerous" for manned aircraft. For this reason, there is an increasing demand for these vehicles in civilian and military applications. In addition,

advancements in unmanned technology allow UAVs to be less expensive, higher performing, and more maneuverable.

1.2 UAV model The ASN-206 shown in figure 1-1 and 1-2 is a twin boom, lightweight, short-range, tactical multi-purpose UAV developed by Xi'an ASN Technology Group Company, a commercial

company owned by NWPU. The ASN-206 can be used for day/night aerial reconnaissance, electronic warfare and countermeasures (EW/ECM), battlefield surveillance, target positioning,

artillery spotting, border patrol, nuclear radiation sampling, aerial photography and prospecting, and electronic countermeasures. It is one of the most popular and advanced tactical UAV systems fielded by the PLA. A versatile aerial platform, the ASN-206 could be fitted with

various mission payloads according to the requirements. The most significant improvement on the ASN-206 is its real-time data link, which transfers video images to the ground control

without any delay, while older-generation UAVs have to be recovered before the photo intelligence can be retreated. Table 1-1 shows the detailed performance of the UAV.


2

Fig.1-1 Original design of the UAV

Table 1-1 Performance of ASN-206 UAV

Specifications Power plant: HS-700 Main airborne

equipment

Wingspan: 6m Length: 3.8m

Height: 1.4m Weight: Max take-off 222kg; Max mission payload 50kg

Speed: Maximum 210km/h Range: 150km

Flight Endurance: 4 - 8 hours Service Ceiling: 5,000 - 6,000m

Total displacement: 692cm3

Compression ratio: theoretical

compression ratio: 10.5 actual

compression ratio: 6.5

Output power: 40kW (54HP) /

6200rpm

Net weight: 22kg

Fuel consumption: 0.51kg / kW.h at

max.power, 0.38kg / kW.h at cruise.

Cooling method: air cooling

Vertical cameras

and panorama

cameras, TV

cameras, infrared

detection

equipment,

positioning and

calibration

equipment and so

on.

The UAV is being widely used by the PLA for various tactical roles. The ASN-206 uses a

tail-pushed, twin-tail braced design. The advantage of this layout is that the propeller driven by

the tail-mounted engine would not disturb the sight of reconnaissance system. The ASN-206 can

be configured with regular or infrared cameras or television seekers which would give the system

a near-real-time capability. The navigation systems of the UAV incorporate GPS and radio


3

command. The original UVA (ASN-206) studied in this thesis has a traditional overall aero

dynamical design [2,3].

Fig.1-2 Rocket take off

1.3 Brief introduction of coaxial counter rotating wing This thesis aims to modify UAV ASN-206 with CCRW, as shown in fig.1-3. The aim is to produce a UAV that will be able to realize self- take off, landing and hovering without significant change of the aerodynamical design of the overall plane. This preliminary design effort will

assist in future development of an UAV which will have the following advantages 1) self-takeoff and landing; 2) hovering capability and high efficiency since there is no tail rotor and therefore

do not need to balance the rotor torque and power consumption; 3) aerodynamic symmetry; 4) vertical and horizontal control efficiency; 5) compact structure, weight and high efficiency; 6) have a greater rate of climb and service ceiling. Therefore rotary wing coaxially reverse rotation

design, the anti-torque of the propeller of the aircraft is not only eliminated, eliminating the tail rotor, and improve the propeller efficiency and stability.

Fig.1-3 UAV ASN-206 with CCRW


4

The rotating wing has no twist angle. It is installed to the suitable angle. Additional engine and control system need to be employed for the UAV. Some coaxial helicopter techniques are involved, but not to the complicated extent. For example, in forward fly, rotating wing only

works before transition period.

Fig.1-4 Self take off

Fig.1-4 Self take off

For cruise condition, the CCRW will be fixed to provide additional lift. For the transition period, the original wing and tail can provide lift, while CCRW decreases and locked in the same

direction of the original wing. Soon after that a pitch control device will be employed to change left top and right bottom wing pitch angle in a small range (5~10°), see fig 1-4. As a consequence, a rotating wing changes to a fixed wing. The USPLSC airfoil is designed

specifically for the CCRW. Because USPLSC airfoil shape is symmetrical to the middle chord position, it requires a smaller range of pitch angle for transition from rotating into a fixed wing

than traditional airfoil (90~180°). Other control system remains same as the basic fixed-wing control system as shown in fig. 1-5.

Fig.1-5 Fixed-wing control system


5

1.3.1 Background and advantages of coaxial rotor In 1859, the British Patent Office awarded the first helicopter patent to Henry Bright for

his coaxial design. From this point, coaxial helicopters developed into fully operational machines

as we know them today. During the first 100 years of aircraft development, fixed-wing aircraft

received most of the attention. Interestingly enough, however, it was helicopter flight that was

first envisioned by man. The Chinese, in ancient times, played with a hand-spun toy that rose

upward when rapidly revolved. This, in fact, marked the first helicopter flight concept[4].

Coaxial rotors are a pair of helicopter rotors mounted one above the other on concentric shafts, with the same axis of rotation, but that turn in contra-rotation. We can see the simple

model of Coaxial Rotor in fig.1-6. The solidity of a coaxial rotor (σ) is defined the same way as for a single rotor:

σ＝

(1.1)

Where b is the total number of blades, c is the blade chord, and R is the radius of the rotor

system. (Note that the disc area used in the above expression is the disc area of just one of the

two rotors, πR2.)[5].

Upper rotor

Lower rotor

Fig. 1-6 Model of coaxial rotor

The advantages of a coaxial design:

One of the problems with any single set of rotor blades is the tendency of the helicopter body to begin spinning in the opposite direction to that of the rotors once airborne. Coaxial rotors solve the problem of angular momentum by turning upper and lower rotors in

opposite directions. However, reactive moments of a coaxial-rotor are compensated for by the counter-rotational forces canceling out each other. This removes the need for any additional

forces like a tail rotor. Without a tail rotor, a helicopter is free from undesirable hazards; there is no more a long tail boom. A shorter helicopter with a smaller footprint can now make the use of a home-based helicopter garage a reality.


6

Furthermore, the reactive moment compensation method employed in a single-rotor helicopter requires the pilot’s constant attention. To achieve balanced flight, the pilot needs to adjust the tail rotor’s side forces; this puts the helicopter to a certain disadvantage compared to a

coaxial design. Given that coaxial rotor helicopters do not have to use power to compensate the reactive moment, coaxial helicopters are generally 16-22% more effective than single-rotor

helicopters. This additional power provides an impressive hover ceiling (500-1,000 m) and vertical rate of climb (by 4-5 m/sec).

Another important feature of the coaxial configuration is illustrated when the helicopter is hovering. The upper rotor airflow grows narrower at 15-20% to the lower rotor and this allows

the lower rotor to suck in additional air. This in turn increases the total rotor airflow and reduces the power used for developing the lift.

No possibility of tail rotor strike; a major cause of helicopter crashes.

Directional stability through cancellation of main rotor gear torque moment (Yaw torque

reaction). Compact size through use of concentric shafts.

Increased pressure differential over rotor system; increased thrust, higher efficiency for

increase in thrust, which translates into a reduction in rotor diameter for a given thrust. The disadvantages of a coaxial design:

Complexity of linkages required to operate pitching control. This disadvantage is

predominantly linked to full-scale aircraft, due to the developments discussed further in the thesis; this is not wholly applicable to SUAV coaxial rotor systems.

Inter-rotor wash interference: reduced efficiency of the lower rotor due to the upper rotor swirling the air in the opposite direction of the lower rotor which requires the lower rotor to run

at higher speed to produce the same lift as the upper rotor. Importance of flow interaction requires for rotor spacing. To ensure sufficiently clean flow

for the lower disc, the spacing must be wide enough to allow as little interaction of the swirl of the upper rotor to impinge on the retreating component of the lower disc[5] .


7

Chapter 2 Aerodynamics of airfoils

In this chapter the aerodynamics characteristics of airfoils are discussed.

2.1 Airfoil The primary lifting surface of an aircraft is its wing. The wing has a finite length called

wing span. If the wing is sliced with a plane parallel to the x-z plane of the aircraft, the

intersection of the wing surfaces with that plane is called an airfoil — having the function of producing a controllable net aerodynamic force by its motion through the air[7]. To be useful this

aerodynamic force must have a lifting component that is much greater than the resistance or drag component. In a powered aircraft, motion through the air is provided by the thrust; so in effect, the airfoil is a device that converts thrust into lift; in a glider the airfoil converts much of the

gravitational force (the potential energy of height) into lift. A typical airfoil section is shown in figure 2-1, where several geometric parameters are illustrated.

Fig.2-1 Airfoil geometric parameters

2.2 Aerodynamic forces The aerodynamics force is the resultant of all forces on a profile in airflow acting on the

center of the pressure. The aerodynamic force has two components – lift which is perpendicular to the relative wind and drag which is parallel to the relative wind. Here the center of pressure is identified. This is the point on which all pressures and all forces act. This point is located where

the cord of a profile intersects with the resultant of the aerodynamic forces lift and drag. This point is expressed as a percentage of the chord of the airfoil. A center of pressure at 30 percent of

a 60-inch chord would be 18 inches aft of the wing’s leading edge. The aerodynamic forces of the lift and drag depend on the combined effect of the many variables- the dynamic pressure the surface area of the profile the shape of the profile and the angle of the attack (The angle at which

the chord c of the airfoil moves in relation to the free stream is known as the angle of attack).


8

F

Fig.2-2 Force vectors on an airfoil

2.2.1 Lift and drag

Lift is the component of aerodynamic force perpendicular to the relative wind. The lifting

force acting on a body in fluid flow can be expressed as

L' = 1/2 ρ v2 A (2.1)

Where L' = lifting force (N), = lifting coefficient, ρ = density of fluid (kg/m3), v = flow velocity (m/s), A = body area (m2)

The lift coefficient is a number that aerodynamicists use to model all of the complex

dependencies of shape, inclination, and some flow conditions on lift. This equation is simply a rearrangement of the lift equation

=

=

where q is dynamic pressure (q=1/2 ρ v2)[7] (2.2)

Here is a way to determine a value for the lift coefficient. In a controlled environment

(wind tunnel) we can set the velocity, density, and area and measure the lift produced. Through division, we arrive at a value for the lift coefficient. We can then predict the lift that will be


9

produced under a different set of velocity, density (altitude), and area conditions using the lift equation.

Drag is the component of aerodynamic force parallel to the relative wind. The drag force acting on a body in fluid flow can be expressed as

D' = 1/2 ρ v2 A (2.3)

Where D' = drag force (N) = drag coefficient.

= D' / (q * A) (2.4)

Lift to drag ratio

The ratio of lift to drag is an indication of the aerodynamic efficiency of the airplane. An airplane has a high L/D ratio if it produces a large amount of lift or a small amount of drag. An

aircraft with a high L/D ratio can carry a large payload, for a long time, over a long distance. Lift and drag coefficients are normally determined experimentally using a wind tunnel. But for some

simple geometry, they can be determined mathematically.

L/D =

⁄ (2.5)

2.2.2 Velocity and pressure distributions

Velocity and pressure are dependent on each other - Bernoulli's equation says that

increasing the velocity decreases the local pressure and vice versa [9]. Thus the upper surface

static pressure in less than ambient pressure, while the lower surface static pressure is higher than

ambient pressure. Ambient pressure is pressure of the surrounding medium, which comes into

contact with the object. This is due higher airspeed (velocity) at upper surface and lower speed at

lower surface of the airfoil (see figure 2-3). It is possible to plot a pressure distribution instead of

the velocity distribution (usually not the pressure, but the ratio of the local pressure to the

stagnation pressure is plotted and called pressure coefficient Cp):

Cp =

(2.6)

Where: P - Static pressure at the point of interest P∞ - Free stream static pressure v∞ - Free stream velocity ρ - Free stream density


10

Fig.2-3 Pressure distribution around an airfoil

The total or stagnation upstream pressure as measured by an impact probe (e.g. a

pitot tube) is the sum of the static and dynamic pressures at that point, i.e.,

= P∞+1/2 ρ v∞ 2 (2.7)

Thus, may also be written in terms of differential pressures:

=

(2.8)

2.3 USPLSC airfoil The geometry of USPLC airfoil is shown in fig 2-4, it is a new airfoil designed with blunt

leading and trailing edge. The geometry of USPLC airfoil is shown in fig 2-4, it is a new airfoil

designed with blunt leading and trailing edge. In a rotating wing, the trailing edge of the retreating blade converts into the leading edge of the fixed wing in forward flight. In order to

accommodate for this hindrance, the airfoil sections of the rotating wing need to have blunt leading and trailing edges. This requirement is fulfilled by employing an USPLC airfoil. However, the blunt trailing edge incurs more profile drag than a sharp trailing edge on a typical

airfoil. Research into techniques for reducing the drag, ranging from active flow control to deployable flaps, continues to be investigated.

The USPLC airfoil is a thin, curved airfoil which can improve the aerodynamic

characteristics in VTOL. The upper surface is protruded and the lower surface is concave. This is

basically to design for the coaxial rotor wing UAV, to get the vertical takeoff and landing (VTOL). The camber distribution and thickness distribution are shown in the fig 2-5 and 2-6.

Maximum thickness of the airfoil is 10% and the max camber is 12% at the 50% of the chord length. The radius on the leading and trailing edge is 1.48/100 of the chord length.


11

V1

V2

,V

3,V

4

0 0.2 0.4 0.6 0.8 1

-0.2

0

0.2

0.4

0.6

0.8

V1

V5

,V

6,V

2

0 0.2 0.4 0.6 0.8 1

-0.2

0

0.2

0.4

0.6

Fig.2-4 Geometry of USPLSC airfoil

Fig.2-5 Camber distribution of USPLSC airfoil

Fig.2-6 Thickness distribution of USPLSC airfoil

As can be seen the both leading and trailing edges are blunt, having the same radius for

both edges .While the maximum camber ratio is almost in the middle of the airfoil. Having the

maximum camber ratio in the middle of the airfoil causes the small lift coefficient, while the drag coefficient is bigger, the blunt trailing edge may bring vortex problem (see fig 2-7).

Both leading and trailing edges are blunt so the stall angle will be smaller, because the

drag will be greater than the other conventional airfoil’s case. To get the cl, cd, cl/cd, and vortex

problem showed by CFD simulations is given in chapter 4. .

Fig.2-7 Vortex structure of USPLSC airfoil at Mach 0.4, AOA = 5°


12

Chapter 3 Airfoil pressure measurement

The experiment helps to obtain pressure distribution on the airfoil and compute the lift and

drag coefficient at different angle of attack at 3 working conditions (20m/s, 30m/s and 40m/s).

3.1 Experiment equipment

3.1.1 Wind tunnel

The experiment is carried out in the LTWT in Northwestern Polytechnical University. It is a low-speed wind tunnel (the air is drawn directly from the surroundings into the wind tunnel and

rejected back into the surroundings), It is 39.52 meters long, structured by steel. There are two available replacement test section: 3D and 2D test section. This paper tests in the 2D test section, the test section dimensions: length × width × height = 3.2 × 1 × 0.4m, Ma = 0.015 ~~ 0.22. . Fig

3-1 shows the components of the wind tunnel.

Fig.3-1 Wind tunnel and test section

3D test section： 1.05×1.2 m，V = 5 ～ 55 m/s

3D test section (3D and 2D serial status) ：1.05×1.2 m，V = 5 ～ 25 m/s

2D test section： 0.4 ×1.0 m， V = 5 ～ 75 m/s

Minimum turbulence level： ε＜ 0.02%

Turbulence level changing range：0.02% ～ 1%


13

3.1.2 Airfoil model

Model size

t=3cm b = 38cm

c=30cm

Fig.3-2 Airfoil model

Model material and pressure taps

The model is made up by paulownia wood. The pressure distribution around the airfoil is obtained from 60 pressure taps (small holes drilled perpendicular to the surface of the blade).

The outer/inner diameter of the copper pipe is 1.2/0.7mm. The pressure taps were placed along the upper and lower surface at the middle of the span in a staggered alignment to minimize disturbances from upstream taps. The taps were drilled directly through the model surface and

into copper tubes lying parallel to the model surface. Pressure tubes (tygon tubes) were connected to the copper tubes and they were lead outside and connected the scanning valve that

sequentially cycles through each pressure tap.

Fig.3-3 Airfoil section shape and pressure taps arrangements


14

3.1.3 Pitot rake

To measure the drag of the airfoil, a pitot rake is employed. It consists of 120 total

pressure probes and 4 static pressure measure tubes, height: 300 mm. It placed back on the airfoil with distance of 0.5 ~ 1 chord lengths to the trailing edge of the airfoil. Pitot rake is perpendicular to the wind tunnel axial direction; it allows the simultaneous measurements of

velocity across the wake.

Fig.3-4 Pitot rake

3.1.4 Airfoil pitch control

The angle of attack of the airfoil in the wind tunnel test section is change using a computer

controlled pitching system that can rotate the airfoil through a full 360 . A motor drives the pitching system , see fig 3-5.

Fig.3-5 Airfoil pitch controller


15

3.1.5 Pressure measuring system

The measuring system employs DSY104 electronic scanning micro pressure measuring

system, manufactured by NWPU. Pressure measuring channel: 192 channels，（±2.5kPa 160

channels，±7.5kPa 32 channels）Canning rate：50000 channel/s.

System precision：±0.1％F.S

3.1.6 Test section

2D test section (see section 3.1.1) is employed to carry out the experiment, is explained.

The setup and wall-mounted airfoil is shown in fig 3-4. 3.2m

Fig.3-4 Test section and wall-mounted airfoil in test section

3.2 Experiment status The wind tunnel experiments are taken at different working condition mention in table

below,

Table 3-1 Working conditions for the experiment

V（m/s） Re Angle of attack (α°)

20 4.11×105 -2,0, 2, 4, 6, 8, 10, 12, 14

30 6.16×105 -2,0, 2, 4, 6, 8, 10, 12, 14

40 8.22×105 -2,0, 2, 4, 6, 8, 10, 12, 14

1.0m

0.4m


16

3.3 Calculation methods

Lift In fig. 3-5 below, the normal pressure force P*ds acting on a small region, may be

resolved into components dY and dX acting perpendicular and parallel to the cord, respectively.

Fig 3-5 Normal pressure force on an airfoil surface

On the upper surface the Y-component is given by:

dY = ds cos where ds=a differential segment of the airfoil surface (3.1)

The normal force Y per unit span and chord is then given by:

dY = dx where dx= ds cos (3.2)

Similarly, Y on the lower surface is given by:

dY = dx (3.3)

Therefore, the total force in Y direction may be written as:

Y = ∫

dx (3.4)

The normal force coefficient then becomes:

=

∫ ( ) ( )

d(x/A) (3.5)


17

Drag

The drag is composed of pressure drag and friction drag. The force coefficient parallel to the chord is responsible for the drag on the airfoil, and this coefficient may be derived in a

manner similar to that previously employed for the normal pressure force. The drag calculated down is only pressure drag. Which is half of the total drag. The total drag is given in reference book[6].

First, divide the airfoil into 2 sections, fore and aft, such that fore is the area in front of the maximum thickness point of the airfoil as indicated in fig.3-6,

Fig.3-6 Pressure regions on an airfoil

For the fore section of the airfoil the x-component is given by:

dX = ds sin (3.6)

The force per unit span acting along the chord line is then given by:

dX = dy where dy=ds sin (3.7)

Similarly, for the aft section the X-component

dX = dy (3.8)

Therefore, the total force per unit span in the x direction can be written as

X = ∫

dy (3.9)

The tangential force coefficient is then given by

∫ ( ) ( )

d(y/A) (3.10)


18

Lift and drag coefficient

Once the numerical values for and have been computed they may be used to

calculate the lift and drag coefficients and using the following force relationships .

= cos - sin (3.11)

= sin + cos (3.12)

Fig.3-7 Component for diagram [10]


19

3.4 Experimental result and analysis The curve of the lift and drag coefficient versus the angle of attack are presented below.

Fig.3-8 at various angles of attack from experimental data

As it can be seen from Figure 3-8, the results show a relationship between the coefficient of lift and angle of attack. When AOA increases from 0° to 6°, increases to maximum value

1.4~1.6. When AOA increases from 6° to 10°, decreases a little bit and increases when AOA increases from 10° to 16°.

Fig 3-9 at various angles of attack from experimental data

Figure 3-9 shows that the minimum is obtained when AOA is in 2 to 4°, the min is

0.02~0.03. increases when AOA increases from 4 to 16°.

0.6

0.8

1

1.2

1.4

1.6

1.8

-4 1 6 11 16

Lift

co

eff

icie

nt

Cl

α

Experiment Cl-α

v40

v30

v20

0.00

0.05

0.10

0.15

0.20

-4 -2 0 2 4 6 8 10 12 14 16 18

α

v40

v30

v20


20

Lift increases proportionately with respect to the angle of attack and wind speed, up to the

critical angle of attack 6°. Lift coefficient do vary with wind speed. These coefficients are useful

to analyze airfoils at different wind speed without performing a wind tunnel analysis for each wind speed.

It can be concluded that the optimum angle of attack is between 4-6°at different velocity

condition. The reason is that at this range the ratio between the coefficient of lift and the angle of

attack is at its maximum. As a result, it is reasonable to assume that in order to obtain maximum lift from USPLSC airfoil, the wing needs to be positioned at 4-6 degrees with respect to the flight

path.

Table 3-2 Experimental , vs AOA results for V=20~40m/s

Value AOA

1.4~1.6 6° ~ 10°

0.02~0.03 2°~ 4°

Stall =1.4~1.6 6° (smooth stall)

Optimum AOA =1.3~1.5,

= 59~79 4°~6°

Pressure distributions curves for the USPLSC airfoil airfoils are shown in chapter 4 with

comparison of CFD.

3.5 Error analysis Every contribution to the flow of the data stream from sensor to reported data is a source

of uncertainty in the final product. In general, the sources of error can be classified as test technique, model, tunnel, instrumentation, and math model related.

Errors of the experiments come from (1) model (2) wind velocity error (3) angle of attack

(4) pressure measuring system. Each one is explained in details:

1. Model

a) Model shape- As it explained before the model is made up by paulownia

wood. When the model is fully made, it slowly contracts when the air is dry and changes shape about some millimeters, which results in error while experiment process.

b) Pressure tap- The error can be caused by introductions of variations in

local tap geometry and/or roughness near the tap hole. Examples of poor design as well as the proper design of the pressure tabs are shown in in fig 3-10. Ideally, to achieve low error, the pressure tap hole need to have a

certain geometry as illustrated in Fig 3-10, (b) with a smooth surface around the hole and perpendicular to the wall. In this case the static

pressure reading is equal to P∞. The assumption can be made that the

streamlines are parallel to the tap surface. In a) lower pressure than P∞ is


21

read as a result; In c) added roughness can cause the flow to slowdown in

the tap region resulting in a pressure reading greater than P∞ In d) have a

similar effect as the roughness.

Fig 3-10 Schematic of pressure tabs. (a), (c) and (d) poor design. (b) proper design

2. Wind velocity error

a) Velocity fluctuation: Both the velocity fluctuation and turbulence of the

wind tunnel will result in error. For low speed experiment, it is difficult to control the velocity at test section to be steady. Because speed is

controlled by the rotating velocity of the blowing fan, it is difficult to maintain the rotating speed at a low value, i.e. any oscillation of the rotating speed will result in a relative wind velocity which cannot be

ignored.

b) Turbulence: Although the LTWT has a small value of turbulence

(Minimum turbulence level： ε＜ 0.02%), it will influences the error. If

turbulence value can be decreased, then error will be decreased.

3. Angle of attacks

An airfoil pitch control shown in fig 3-5 can cause error. Motor drives inside an airfoil pitch controller device change the AOA. While

changing AOA it doesn’t give a high accurate AOA. For example AOA 8° it can change the airfoil to AOA 7.8° or 8.1°.

4. Pressure measuring system

DSY104 electronic scanning micro pressure measuring system is

employed to measure pressure distribution which has a certain level of error.


22

Chapter 4 CFD simulations

This simulation of flow passing the airfoil is performed, using computational fluid dynamics (CFD) software Fluent. CFD results of the surface pressure distribution, lift and drag

force acting on the airfoil are compared with experimental data. The simulation been has developed after the laboratory experiment and therefore we use the same USPLSC airfoil

geometry as well as the same flow parameters, such as fluid density, viscosity, angle of attack, and free stream velocity.

CFD is the analysis of systems involving fluid flow, heat transfer and associated phenomena such as chemical reactions by means of computer-based simulation. The technique is

very powerful and spans a wide range of industrial and non- industrial application areas. The main reason why CFD has become so popular is the availability of affordable high performance computing hardware and the introduction of user-friendly interfaces have led to a recent upsurge

of interest, and CFD has entered into the wider industrial community since the 1990s. CFD codes are structured around the numerical algorithms that can tackle fluid flow problems. In order to

provide easy access to their solving power all commercial CFD packages include sophisticated user interfaces to input problem parameters and to examine the results. Hence all codes contain three main elements: (i) a pre-processor, (ii) a solver and (iii) a post-processor.

In the fluid dynamics, there are many commercial CFD packages available for modeling

flow in or around objects. The computer simulations show features and details that are difficult, expensive or impossible to measure or visualize experimentally. The simulation is performed on the commercial CFD code Fluent. Fig 4.1 shows the structure of Fluent software, where pre-

processing is completed in software. Gambit is a mesh generator. First, Gambit is used to make a discretization of flow domain, and then Fluent is applied to solve the flow[11].

Fig.4-1 The composition of the CFD solver


23

4.1 Gambit and Fluent Gambit is the tool to get the layout for the model and then generate the grid. We get the

mesh for model by Gambit and then use the mesh into Fluent to get our desired results. Gambit,

like similar programs such as ANSYS, constructs its geometry by references to a hierarchy of geometric "Entities". Entities have to be set up in the order: vertices - edges - faces – volumes. Vertices are points defined by three coordinates. Edges (straight lines, circles, curves, etc.) are

constructed by reference to vertices. Faces (flat and curved) are constructed from Edges. Volumes are formed by stitching Faces together (but only face that share common edges). Two-

dimensional modeling only goes as far as a set of planar Faces. You must start with good design sketch, showing in particular the vertices (with their coordinates) and the edges. You also need to decide how you will split your model into Volumes, because the best meshing schemes will

need special attention to the Volume structure.

There are many types of grid we can generate by Gambit, such as, Triangle, quadrilateral, tetrahedron, hexahedron, pyramid, prism/ wedge. Fig4-2 below shows the geometry of these types of this grid [10]. Quadrilateral grid is also known as local mesh, which we are also using in

our case to get CFD simulation of USPLSC airfoil. Geometry and grid are saved in a database file (*.dbs) and mesh is saved into a solver-dependent file (*.msh)

Fig.4-2 Various grid in Gambit

Fluent solver is based on finite volume method; it can handle both structured grids, i.e.

rectangular grids with clearly defined node indices, and unstructured grids. Unstructured grids are generally of triangular nature, but can also be rectangular. In 3-D problems, unstructured grids can consist of tetrahedral (pyramid shape), rectangular boxes, prisms, etc. "Fluent" is the

general name for the collection of CFD programs sold by Fluent, Inc. of Lebanon. Domain is discretized onto a finite set of control volumes (or cells). General conservation (transport)

equations (see eq. 3-1) for mass, momentum, energy, species, etc. are solved on this set of control volumes. Partial differential equations are discretized into a system of algebraic equations. All algebraic equations are then solved numerically to render the solution field. (3.1)


24

Today, thousands of companies throughout the world benefit from the use of ANSYS Fluent software as an integral part of the design and optimization phases of their product development. Advanced solver technology provides fast, accurate CFD results, flexible moving

and deforming meshes, and superior parallel scalability. User-defined functions allow the implementation of new user models and the extensive customization of existing ones. The

interactive solver setup, solution and post-processing capabilities of ANSYS Fluent make it easy to pause a calculation, examine results with integrated post-processing, change any setting, and then continue the calculation within a single application.

Table 4-1 The general sequence of operations for Fluent; FLUENT: getting started

1. <Fluent>; <2> when asked for version "3" means a 3D model

2. File - Read - Case – channel.msh Remember that Fluent doesn't carry the model

geometry data. Geometry and mesh changes have to

be done in Gambit

3. Grid - Scale

Make sure the grid is the size

you expected, and apply scale

factors

By default, FLUENT assumes lengths are in metres.

GAMBIT doesn't have units. You have to ensure that

Fluent knows which units you were using (mm in this

case).

4. File - Write - Case

(later, when you have results, select Case &

Data)

Do this from time to time. The Case file

contains your fluid model and mesh.

5. Display - Grid Check that it's the right model. Learn how to

pan and zoom the display.

6. File - Hardcopy - select graphics format -

Save

Saves the active graphics window

7. Surface - Lines Enter the two sets of

coordinates (100,0); (100,5); to define a line

half way along the pipe

Define any planes, lines or points over which

you might want to display information.

8. Define - Models - Viscous Default viscous model is Laminar-note the

many alternatives

9. Define - Materials - pick or define a fluid The default fluid is "air"

10. Define - Boundary conditions

Set "inlet" to Velocity Inlet; enter velocity

of x m/s

Set "outlet" to Outflow

Set “top” and “bottom” to Wall

By default, FLUENT assumes velocities are in

m/s. (Note second chance to give a ‘type’ to

the bc’s. But only if they are separately

named.)

11. Solve - Initialise - Init Otherwise the computer array is full of the

junk from the last user.


25

12. Solve - Monitors - Residual - Plot Displays progress of iteration. Note options

available

13. Solve - Iterate - set 100 iterations Solution stops automatically at 100 or when

residuals reach 10-3.

14. Display – Vectors of velocity – Contours of

velocity - Contours of pressure

Look at results on selected planes

15. Plot-Velocity Draws graphs of data along selected lines

16. Define – Boundary conditions

Change “inlet” velocity

Solve-Iterate

Adjust the inlet velocity until the velocity

profile is just fully formed at the outlet. Not

necessary to re-initialise before solution.

17. Plot - Velocity Draw the graph of the velocity profile at the

mid-way point and save this

18. Solution – 2nd Order –Solve Change the order of the solution to 2nd order

(you have been working to 1st order). Again

not necessary to initialise.

19. Plot – Velocity Compare the velocity profile at the mid-way

point with that obtained using the 1st order

solution.

Case set-up and boundary conditions are saved in a case file (*.cas) and solution is saved in a

data file (*.dat)[12,13].

4.2 CFD simulation of USPLSC airfoil CFD simulation of USPLSC airfoil is done by Gambit and Fluent. Mesh is generated by

software Gridgen then imported to Gambit. The mesh contains 80,000 cells, circulation division

× radial division is 400×200. Then use the Fluent solver is used to get results for USPLSC airfoil at various angle of attack and Mach numbers.

The global mesh for USPLSC airfoil is shown in fig 4-3 below. Fig 4-4 below shows the

local mesh distribution for USPLSC airfoil.


26

Fig.4-3 Global mesh distribution of USPLSC airfoil

Fig.4-4 Local mesh distribution for USPLSC airfoil


27

4.2.1 Establishment of CFD model Boundary condition in this model is pressure far field, fluid shown in the fig 4-5 below.

Fig.4-5 Boundary conditions setting in Fluent

Turbulent flow model realizable k-e is selected, see fig 4-6.

Fig.4-6 Viscosity setting in Fluent

For different velocity change Mach number, for different AOA change X-component of flow direction/Y-component of flow direction. When solution is done, , can be obtain for

different working conditions.


28

4.2.2 Working conditions

As, mentioned in chapter 3 section 3.2(experiment status), the CFD simulation is

completed over velocity 20 to 40m/s. Next section 4.2.3 below show the some graph along with experiment result. Section 4.2.4 only represents the graph at velocity 20m/s.

4.2.3 Comparison of ,

Fig.4-7 , varying AOA at v=20m/s



29


From these three graphs it can be observed that the agreement between experiment and

CFD is satisfied.

The comparison between experiment and CFD is good above 7.5° AOA.

Below 7.5° AOA the k-e prediction is maximum difference of .

The CFD error of is -27~-6% compared with experimental result (for v=20m/s).

4.2.4 Comparison of Cp Comparison of pressure distributions at various AOA for the USPLSC airfoil is shown

below.

There is some scatter in the CP values at the top of the airfoil near leading edge region until

x/c = 0.17, caused by either uncertainty of the calibration or small irregularities of the surface or

the pressure taps.


30

Fig.4-10 cp comparison at AOA=00


-1.5

-1

-0.5

0

0.5

1

1.5

-0.2 0 0.2 0.4 0.6 0.8 1 1.2

CFD Cp result

Exp Cp upper

Exp Cp lower

-2

-1.5

-1

-0.5

0

0.5

1

1.5

-0.2 0 0.2 0.4 0.6 0.8 1 1.2

CFD Cp result

Exp Cp upper

Exp Cp lower

Cp

Cp

x/c

x/c


31



-2

-1.5

-1

-0.5

0

0.5

1

1.5

-0.2 0 0.2 0.4 0.6 0.8 1 1.2CFD Cp result

Exp Cp upper

Exp Cp lower

-3.5

-3

-2.5

-2

-1.5

-1

-0.5

0

0.5

1

1.5

-0.2 0 0.2 0.4 0.6 0.8 1 1.2

CFD Cp result

Exp Cp upper

Exp Cp lower

Cp

Cp

x/c

x/c


32

As can be seen from 4 graphs when AOA increases Cp difference between experiment and

CFD decreases. This is the reason why difference is small at high AOA between experiment

and CFD than small AOA.

In order to understand the static pressure distribution and streamline across the airfoil, figs

4-14 to 4-17 present the results from CFD calculations. As can been seen from figs 4-14 to 4-17,

there is a flow separation at the rear of the upper surface. The pressure distribution of the airfoil

is high pressure at lower surface and low pressure at upper surface.

Fig.4-14 Static pressure distribution(Pa) and streamline for AOA=0 at v=20m/s



33




34

Chapter 5 Test-rig design for coaxial rotor

The aim of this chapter is to design a test-rig. The test-rig priority is to be able to test and

measure coaxial fixed pitch system configuration variables in hover. The rotor test stand consists

of a test stand, rotating device, rotor shaft, rotor hub, instrumentation for sensors, and a data acquisition system. The simple organization of rotor test stand is below in fig.5.1.

Fig.5-1 Organization of rotor test stand

The investigation of the coaxial rotor system will primarily revolve around these following

testing variables:

Inter-rotor spacing – The space between upper rotor and lower rotor. Inter-rotor spacing

is one of the fundamental components of the UAV coaxial system which has been tested due to the associated aerodynamic effects; interference-induced power losses, wake

contractions, and rotors vena contracta[14]. The H/D ratio is used as a non-dimensional figure to enable comparison of multiple systems across a range of scales. The H/D ratio is given as:

(5.1)

Fig.5-2 Inter-rotor separation distance


35

The different research and experiment has different point of view on H/D ratio. Some of them are mention below, Table 5-1 Different thought about H/D ratio [15]

The table below demonstrates that the SUAV example systems have a significantly higher H/D ratio (average H/D=0.315), when compared with the average full-scale helicopter systems having an H/D=0.09. In our experiment we are going to start H/D ratio

from 0.1.

Fig.5-3 H/D ratio comparison chart [16]

Pitch Control- For maximum efficiency we need to control the angle of attack. Now for a coaxial rotating system the lower rotor is meeting wash down air. In general the upper

rotor will cause a "swirl" of air behind it. The air behind the upper lower is travelling backwards faster now, but it also has a rotation applied to it. The increased backwards speed of the airflow tends to require a higher pitch on the lower rotor; the rotation tends

to cause a requirement for a lower pitch. The lower rotor has different pitch requirements to the upper, however just what those requirements are changes depending on the type,

speed, use etc. of the system.

Bohorquez( MAV) Syal He noted that “rotor spacing has a

limited effect on the coaxial rotor performance and is not a critical parameter that has a dramatic effect

on performance”

Syal remarks that “higher inter-rotor spacing is

desired to reduce the induced losses of the coaxial rotor system in hover. With a higher inter-rotor spacing, a small fraction of the lower rotor lies in

the slipstream wake generated by the upper rotor”

H/D ratio =0.714 the spacing recommendation is given at 75% of the rotor radius.


36

We are going to test on fixed pitch. Once rig is shutdown we manually change the

pitch and run the rig.

Thrust- Thrust is composed of two different rotors. A precise measurement should be

able to obtain the thrust of the upper/lower rotor separately. But the literature review of the coaxial test-rig shows most of research contains only one thrust device. Thrust is

measured with varying collective pitch angle at fixed RPM. Most of the reference uses the load cell to measure thrust. We will also use thrust load cell. Torque also can be measure by load cell.

Power/torque- we are going to use torque device（strain torquemeter） to measure

torque.

Power(w) = torque(N*m)*Angular velocity (rad/s)

Rotational speed- There is different way to measure rotational speed. Many lab use photo sensors to measure the speed. The infrared speed detector is used in our lab to

measure rotational speed.

5.1 Test-rig development We have two options to construct the test-rig. Option 1 Test-rig of face to face rotor where

inter-rotor space is manually varied and option 2 test-rig of serial rotor. The different prototype test-rig deigns and their working condition is explained below.

Option 1.Test-rig of face-to-face rotor [17] The components used in the setup for a coaxial rotor system (using HALO’s components

as a datum) have dictated the majority of the test-rigs overall design.

The motors used for the coaxial rotor system are the AXI 2217/20 electric out runner DC

motors.

Linear motion technology in the form of a motor driven lead-screw system was chosen to

vary the inter-rotor spacing of the coaxial rotor configuration.

For the measurement of thrust, the Autonomous System Lab’s thrust testing rig is

incorporate, which works on a fulcrum lever principal.

One of the foundation attributes for the coaxial rotor systems experiments is to be able to

test if the configuration is in the condition of equal torque, the test-rig will incorporate a yaw

effect measurement system.


37

Fig. 5-3 CAD model of test-rig of face-to-face rotor [16]

Option 2.Test-rig of serial rotor[18] The rotor test stand was constructed and integrated with the various components, as shown

in Fig.5-4.

The rotating device is composed of a rotor hub and a power transmission. The power transmission equipment consists of BLDC (BrushLess Direct Current) motor

and the electric power is supply through a 5 KW rectifier.

LabView 7.0 was used for signal conditioning and data acquisition. Thrust and torque is measured by load cells.


38

Fig. 5-4 Thrust and torque measurement

In addition, the instrumentation includes photo sensors for measuring the rotational speed, sensors for measuring the voltage and current, and a 3-axis accelerometer for measuring the

vibration of the rotor test stand post.

Fig. 5-5 Photo sensor and accelerometer

The data measured by each sensor are confirmed and gathered at a control PC through the data acquisition system in real time, as show in fig.5-1 The rotational speed and the collective pitch angle of the rotors are controlled by the control PC through an Electronic Speed Controller,

a servomotor, and a servo actuator .Tests of the hover performance of the rotor blade were performed in a way that minimized the ground effect and confirmed the reliability of the

experiments at a test station in the Korea Aerospace Research Institute. The rotor test stand was constructed and integrated with the various components, as shown in Fig. 5-6.


39

Fig.5-6 Test-rig of serial rotor [18]

5.2 Single rotor test-rig at NWPU We have the single rotor test-rig in our lab. See the fig. 5.1 to understand the working

condition. This is the detail picture of the test-rig. This is composed of three main parts- base

part, support part and thrust measure part.

Base part is consisting of rotating device. It’s fixed on the ground safely.

Support part connects the base part with thrust measure part. In base area the motor is

connect with spline shaft with cardan joint to thrust measure part. The reason to connect the shaft

with cardan joint is because the thrust measure part should be free movement on the vertical

direction and it allow for variations in the alignment and distance between the drive shafts

frequently.

Thrust measure part is connecting to the rotor hub. Load cell is placed on the platform of

the thrust measure part.


40

Fig.5-7 Single Rotor Test-rig at NWPU

A-Thrust measuring part

B- Support part

C- Base part

D-Thrust measuring device

E- RPM measuring device

F- Shaft

G-Motor

5-3 Development of coaxial rotor test-rig

1. Face gear-drive for coaxial propellers – The face gear is a circular disc with

a ring of teeth cut in its side face; hence the name faces gear. Both an inner and

outer shaft is spirally threaded. The inner threaded shaft is thinner in diameter and

is residing in then outer shaft is greater in diameter. As shown in the fig below.


41

Fig.5-8 Face gear-drive for coaxial propellers

Face gear drive has the following advantages:

• Axial freedom of the pinion. This means:

- No need for the exact axial positioning of the pinion

- Tolerable contact pattern changing

• The pinion is a cheap, easy to manufacture spur gear

• No forces in the direction of the pinion axis

• Generous tolerance of the distance between the pinion axis and the head plane of

the face gear.

2. H/D ratio - The rotor hub is fixed in shaft with help of nut and lock washer.

Here the advantage of this configuration is that to calculate H/D ratio between the

rotor hubs. It can be fixed at variable height as required, by locking them

respectively with lock washers and nuts. Shown in the figure below.

Fig.5-9 Rotor hub, nuts, lock washers


42

3. Pitch control-In this manual configuration for changing the pitch of the blades

nuts and lock washers are fasten the blade grip from inside as well as outside.

Whenever it needs to change the pitch, it can be simply loosen the outside nut,

change the pitch angle and then fasten it again to fix it in the desired angle. The

advantage for this setting is that it’s simple configuration as well as cheaper in

cost to manufacture.

Fig.5-10 Rotor hub, Nuts, Lock washers, Blade grip, Horizontal hinge pin

Fig.5-11 Coaxial rotor test-rig


43

5.4 Safety devices The rotors of the test-rig reach at high revolution speeds. This means the test-rig is operated beyond the first resonance speed. Considering the rotor's weight of several hundred

kilograms and the respective inertia a tremendous amount of energy stored in the rotors during operation. In order to handle this situation two kinds of precautions against damage are taken.

First of all the entire test-rig is comprehensively monitored during operation. This condition monitoring includes sensors for vibrations and rotation speeds as well as sensors

monitoring pressures and flow rates of the lubrication system. Other sensors monitor the state of the propulsion system. All these data are evaluated in a programmable logic control that surveys the test-rig and can initiate an emergency shut down if necessary.

On the other hand the test-rig is built in a steal net. In case of a disaster the net will contain

the debris and save the surrounding structures. The operation of the test-rig is done from a surveillance room separated from the test-rig.

net

Test-rig

Fig.5-12 Test-rig enclosed by net


44

Chapter 6 Conclusion and future work

6.1 Conclusion In this thesis it has been studied to improve in UAV technology, primarily the CCRW

UAV concept, resulted in the conclusion that a parametric study into the aerodynamic performance of fre-and-aft symmetrical USPLSC airfoils would benefit future research in CCRW technology development and could serve as a tool for new coaxial rotor-wing design

concepts.

Wind tunnel tests with approximately 2D flow were carried out for the USPLSC airfoil in the LTWT wind tunnel binary test section at velocity 20~40m/s to obtain the airfoil surface pressure distribution data. To predict the aerodynamic performance at the each part of the wing,

3 different working conditions are employed. Each condition has one precise velocity and 9 different AOA. The result obtains from wind tunnel showed that the airfoil behaved well

according to the design assumptions. In order to obtain maximum lift from USPLSC airfoil, the wing needs to be positioned at 4-6º with respect to the flight path. The curve of the lift coefficient versus the angle of attack shows a stall at 6º. The stall is smooth.

CFD simulation has performed over 2D USPLSC airfoil. Comparisons between experiment and CFD simulation result were carried out with the USPLSC airfoil. The

comparisons were in good agreement with their results. The result shows when AOA increase

the Cp difference between experiment and CFD decrease. This is the reason why , difference is small at high AOA between Experiment and CFD. The CFD error of cl is -27~-6% compared

with experiment result for v=20m/s. Conceptual design of the test-rig is illustrated in last chapter. The coaxial gear box is

design on the basis of face gear mechanism. The design features manually changeable the inter-rotor spacing and pitch.

6.2 Future work To understand the more detail about the airfoil, it need to perform high speed

wind tunnel experiment 50~100m/s. In the main, refinement of model and measurement technology improvements in wind tunnel.

Investigation on CFD error analysis and improvement in the development of airfoil model. A more complete grid convergence study and a further validation

study would both need to be conducted to develop better techniques and methodology for aerodynamic USPLSC analysis of airfoils.

This thesis presents the conceptual design of test-rig which capable of measuring the thrust and torque of a rotor blade in the hover state. In order to construct more

detail should look after structure design, power supply and control technics.

Ongoing and future work will continue to investigate the feasibility of aerodynamics

performance based in real life environments through both analysis and testing. Key emphasis


45

will be placed on investigating solutions that fit into the current design and operational paradigms, as well as looking at retrofittable solutions.


46

References

1. http://www.thefreelibrary.com/The+UAV+as+sensor+platform+--+from+pioneer+to+global+hawk%3A+the...-a080534446

2. http://www.sinodefence.com/airforce/uav/asn206.asp

3. http://www.tjskl.org.cn/suppliers/czae09dc/h-xi_an_asn_technology_group_co_ltd.html

4. J. Gordon Leishman. Principles of Helicopter Aerodynamics. Cambridge University Press, New

York, 2000.

5. Colin P. Coleman. A survey of theoretical and experimental coaxial rotor aerodynamic research. Ames Research Center, Moffett Field, California, 1997.

6. http://www.gyrodynehelicopters.com/coaxial_benefits.htm

7. John D . Anderson Jr. Fundamentals of aerodynamics. McGraw-Hill Education, 2005.

8. Dava Newman. Interactive aerospace engineering and design. Massachusetts Institute of

Technology, McGraw-Hill Education, 2002.

9. http://www.princeton.edu/~asmits/Bicycle_web/Bernoulli.html

10. M R Emami. Aerodynamic forces on an airfoil. AER 303F, Aerospace Laboratory I, University of Toronto, 2007.

11. H K Versteeg and W Malalasekera, An introduction to computational fluid dynamics. Pearson

Education Limited, England, 2007.

12. FLUENT 6.3 User Guide, FLUENT Inc. of American. 2006

13. GAMBIT 6.2 User Guide, FLUENT Inc. of American. 2006.

14. J.C. Bell; M. Brazinskas; S. D. Prior; C. Barlow; M. A. Erbil; M. Karamanoglu. Development of a Test-Rig for Exploring Optimal Conditions of Small Unmanned Aerial Vehicle Co-Axial Rotor Systems. Department of Product Design and Engineering, School of Engineering and Information Sciences, Middlesex University, Trent Park Campus, Bramley Road, London, N14 4YZ, UK

15. J.C. Bell; M. Brazinskas; S. D. Prior; C. Barlow; M. A. Erbil; M. Karamanoglu. Development of a

Test-Rig for Exploring Optimal Conditions of Small Unmanned Aerial Vehicle Co-Axial Rotor Systems. Department of Product Design and Engineering, School of Engineering and Information Sciences, Middlesex University, Trent Park Campus, Bramley Road, London, N14 4YZ, UK

16. J. Bell. Investigations into Optimal Co-Axial Rotor System Configurations for Small UAVs

Master’s Thesis, Middlesex University, UK, 2010.


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17. Stephen D. Prior and Jonathon C. Bell. Empirical Measurements of Small Unmanned Aerial

Vehicle Co-Axial Rotor Systems. Journal of Science and Innovation, Vol. 1, No. 1, Middlesex University, London, UK. 2011.

18. Byoung-Eon Lee, Young-Seop Byun, Jeong Kim and Beom-Soo Kang. Experimental hover

performance evaluation on a small-scale rotor using a rotor test stand. Journal of Mechanical Science and Technology 25 (6) (2011) 1449~1456, Pusan National University, Busan, Korea. 2011.


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Acknowledgment I express my deepest gratitude to my thesis advisor Associate Professor Zhao Xu for

making this work possible. Associate Professor Zhao Xu, with her unique style of guidance saw

me through every aspect of this work and kept me focused to finish this work, while inspiring me to work on many other topics. I would like to thank her for the freedom she allowed in the course of this investigation. It is a pleasure and honor being her student.

I have the greatest sense of appreciation to my friends and colleagues for making past four years so enjoyable.

Lastly I would like to acknowledge my family for their love and encouragement during

this process. They have been very helpful from time to time. Specifically my father for his persistence in pushing me towards my goal of completing this thesis.


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Graduation project summary

I attended the NWPU for several years to earn my degree. In four years of university is

coming to an end, learn many things and have really good time. At the end of this 4th year, I am

writing my Undergraduate thesis. Undergraduate thesis is considered as a significant task in the

life of a student. With this experience, there are lots of knowledge and lessons that i have learn

from it. Moreover, with the undergraduate thesis as the first experience for writing technical

paper, it will require a the knowledge I have learned in past four year in university, lot of time

and effort, for which I am very thankful to my supervisor Zhao Xu. Half of this semester I

continue to learn, accumulate and improve my knowledge. Find many useful information from

study material and internet, getting familiar with physical theories and communicate with other

students, and most importantly with my teacher’s guidance, I am able to complete the graduation

thesis. During this time I have learned and understand many new things and get familiar with

Experimental research on aerodynamics performance of an airfoil and test-rig design of coaxial

rotating wings. This thesis is a valuable learning experience for me.

Documents

khagendra thesis