AERODYNAMIC DATA GENERATION AND

DESIGN SUPPORT FOR SOLAR UAV:

WIND TUNNEL TESTING

Submission Report for Undergraduate Award 2014

Abstract The Solar Unmanned Aerial Vehicle (Solar UAV) project aims to design and fabricate a medium

altitude long endurance (MALE) UAV with 24-hour continuous flight mission using solar energy

as the sole power source. The aircraft has a high wing and T-tail configuration, with a wing span

of 17.7m and chord length of 1.26m. Having a target weight of 85kg, the aircraft is targeted to fly

at 8000m altitude with flight speed of 14m/s. To have a better understanding of the flight

characteristic for further design evaluation of the UAV, a 1:9 scale wind tunnel model was

designed to generate aerodynamic data of the aircraft.

The wing and horizontal tail were made out of carbon fibre reinforced plastic (CFRP) and foam

core structure. The other major components of the model were made out of aluminium and

stainless steel. 3D-printed components like control surfaces, wing tips etc were also incorporated

in the model. The model weighs only 5.96kg and is well within the desirable weight category for

the balance chosen, meeting measuring accuracy. The model also meets the safety requirement

W H NAA WS TW W CW; (1) and low speed wind tunnel practice of limiting the model span to 80% of tunnel width (2) and flow blockage below 3-5%.

The wind tunnel test was carried out in the 3m x 2.25m low speed wind tunnel of National Wind

Tunnel Facility at Indian Institute of Technology Kanpur to generate aerodynamic force and

moment data at test speed of 63.4m/s matching the flight Reynolds number of 610000. The

aircraft was found to have good lateral and directional stability and better lifting characteristics

than the empirical estimation. However, the aircraft exhibits longitudinal instability and this

requires second tunnel entry for test results verification and further aircraft design

improvements to enhance the stability of the aircraft.

Table of Contents Abstract ............................................................................................................................................. I

List of Figures ..................................................................................................................................... I

List of Tables .................................................................................................................................... IV

List of Symbols .................................................................................................................................. V

1. Introduction .............................................................................................................................. 1

1.1. Project Overview .............................................................................................................. 1

1.2. Project Scope .................................................................................................................... 2

2. Literature Review ..................................................................................................................... 3

2.1. Model Design and Accuracy ............................................................................................. 3

2.2. General Guidelines for Wind Tunnel Tests ....................................................................... 4

2.2.1. General test procedure ............................................................................................ 4

2.2.2. Permissible Measuring Errors of the Load Balance .................................................. 5

3. National Wind Tunnel Facilities ................................................................................................ 6

3.1. Balance Selection and Balance Mounting Scheme........................................................... 7

4. 1:9 Scale Wind Tunnel Model Design ....................................................................................... 9

4.1. Design Evolution ............................................................................................................... 9

4.2. Model Component Description ...................................................................................... 11

5. Model Stress Analysis ............................................................................................................. 19

5.1. Mechanical Properties of Materials and Aerodynamics Inputs on Component Loads .. 19

5.2. Factor of Safety ............................................................................................................... 19

5.3. Stress Analysis of Model Structural Components .......................................................... 19

5.3.1. Wing ........................................................................................................................ 19

5.3.2. Horizontal Tail ......................................................................................................... 21

5.3.3. Tail boom ................................................................................................................ 21

5.4. Analysis of the Structural Joints ..................................................................................... 24

5.4.1. Wing Joint ............................................................................................................... 24

5.4.2. Wing Mounting Block Fasteners Analysis ............................................................... 26

5.4.3. Horizontal Tail Joint Analysis .................................................................................. 27

5.4.4. Tail Mounting Block Fasteners Analysis ................................................................. 28

5.4.5. Vertical Fin Joint Analysis ....................................................................................... 28

5.4.6. VT Empennage Loft Fastener Analysis .................................................................... 29

5.4.7. Tail Boom Joint Analysis ......................................................................................... 29

5.5. Thread Engagement Length ........................................................................................... 30

5.6. Wing and Tail deflection ................................................................................................. 30

5.7. Summary ......................................................................................................................... 33

6. Hinge Moment Estimation and Servo Selection ..................................................................... 34

6.1. XFOIL Computation:........................................................................................................ 34

6.2. EI; A;I C W; ............................................................................. 36

6.3. Results: ........................................................................................................................... 37

7. Manufacturing Process of the Wind Tunnel Model ............................................................... 39

8. Model Assembly and Inspection ............................................................................................ 41

9. Control Surface Actuation System and Calibration Process ................................................... 44

9.1. Control Calibration Process ............................................................................................ 44

9.2. Control Surface Loading Test .......................................................................................... 45

10. Wind Tunnel Test and Data Analysis .................................................................................. 48

10.1. Wind Tunnel Test Matrix ............................................................................................ 48

10.2. Wind Tunnel Test Procedures .................................................................................... 48

10.3. Measurement Accuracy .............................................................................................. 51

10.4. Data Analysis .............................................................................................................. 52

10.4.1. Transformation Matrix ........................................................................................... 52

10.4.2. Velocity Sweep Test ................................................................................................ 53

10.4.3. Step vs Sweep Test ................................................................................................. 54

11. Conclusion .......................................................................................................................... 57

12. References .......................................................................................................................... 58

Appendix A ..................................................................................................................................... 60

Appendix B ...................................................................................................................................... 62

Appendix C ...................................................................................................................................... 87

Appendix D ..................................................................................................................................... 91

Appendix E ...................................................................................................................................... 97

Appendix F .................................................................................................................................... 102

Appendix G ................................................................................................................................... 104

Appendix H ................................................................................................................................... 106

I

List of Figures Figure 1: Mission Profile ................................................................................................................... 1

Figure 2: Half Model Test (1) Figure 3: 2D Wing Test ................................................................ 3

Figure 4: Fibre-Glass Model (4) Figure 5: Laminated Mahogany Model NA-73X (5) ................ 4 Figure 6: DAQ System ....................................................................................................................... 6

Figure 7: Balance G ........................................................................................................................... 7

Figure 8: Balance G Assembly Figure 9: Balance Interface Unit .................................................. 8

Figure 10: Balance G Setup ............................................................................................................... 8

Figure 11: Conceptual Model ........................................................................................................... 9

Figure 12: Final Design.................................................................................................................... 10

Figure 13: Fuselage Cap Figure 14: Fuselage Cap (3-view Drawing Extracted) ......................... 11

Figure 15: Fuselage Figure 16: Fuselage (3-view Drawing Extracted) ................................... 12

Figure 17: Wing Mounting Block Figure 18: Wing Mounting Block (3-view Drawing Extracted)

................................................................................................ 12

Figure 19: CFRP Foam Core Composite Wing ................................................................................. 12

Figure 20: Wing Bottom View ........................................................................................................ 13

Figure 21: Tailboom Connector Figure 22: Tailboom Connector (3-view Drawing Extracted) .. 13

Figure 23: Fuselage Patch Figure 24: Fuselage Patch (3-view Drawing Extracted) .................. 14

Figure 25: Tailboom ........................................................................................................................ 14

Figure 26: Empennage III Figure 27: Empennage III (3-view Drawing Extracted) .................... 14

Figure 28: VT Empennage Loft Figure 29: VT Empennage Loft (3-view Drawing Extracted) .... 15

Figure 30: Vertical Fin ..................................................................................................................... 15

Figure 31: Tail Mounting Block Figure 32: Tail Mounting Block (Side View) ........................... 16

Figure 33: Horizontal Tail ................................................................................................................ 16

Figure 34: Horizontal Tail Bottom View.......................................................................................... 16

Figure 35: Rocker Arm Assembly 1 Figure 36: Rocker Arm Assembly 2 .................................... 16

Figure 37: Wing Cap Figure 38: Horizontal Tail Cap ................................................................. 17

Figure 39: Elevator .......................................................................................................................... 17

Figure 40: Levelling Block and Model Stand Figure 41: Model On The Stand ............................ 18

Figure 42: Wing Bending Moment ................................................................................................. 19

Figure 43: Airfoil Approximation 1 ................................................................................................. 20

Figure 44: Tail Boom Stress Analysis .............................................................................................. 22

Figure 45: Tail Boom Cross Section ................................................................................................ 23

Figure 46: Tail Boom Torsion Analysis. ........................................................................................... 23

Figure 47: Wing Joint Illustration ................................................................................................... 24

Figure 48: Tensile Load Due To Bending And Rolling Moment Figure 49: Tensile Load Due To Lift

.................................................. 25

Figure 50: Tensile Load Due To Pitching Moment .......................................................................... 25

Figure 51: Wing Mounting Block M5 Fasteners Figure 52: Tensile Load Due To Pitching Moment

......................................................................... 26

Figure 53: Tensile Load Due To Rolling Moment ............................................................................ 26

Figure 54: Tail Joint Analysis Figure 55: HT Lift Bending .......................................................... 27

II

Figure 56: Tail Mounting Block Fastener Analysis .......................................................................... 28

Figure 57: Fin Joint Analysis............................................................................................................ 28

Figure 58: VT Empennage Loft Fastener Analysis........................................................................... 29

Figure 59: Tail Boom Joint Analysis Figure 60: M4 Fastener Internal Load ............................ 29

Figure 61: Wing Lift Distribution .................................................................................................... 31

Figure 62: Tail Boom Upward Bending Deflection ......................................................................... 32

Figure 63: FBD For Tail Deflection Analysis .................................................................................... 32

Figure 64: Tail Boom Defection Due To Side Force ........................................................................ 32

Figure 65: GDES Function ............................................................................................................... 34

Figure 66: Experimental Data of() (11) ............................................................................... 35 Figure 67: Ch vs deflection angle (sigma) ....................................................................................... 37

Figure 68: Ch vs deflection angle (empirical) ................................................................................. 37

Figure 69: JR DS181 Slim Wing Servo (Courtesy of hobbyking.com (12)) ...................................... 38

Figure 70: Metallic Components .................................................................................................... 39

Figure 71: CFRP Wing and Tail ........................................................................................................ 40

Figure 72: Wing Load Test .............................................................................................................. 40

Figure 73: Model Full Assembly Front View ................................................................................... 41

Figure 74: Model Full Assembly Side View ..................................................................................... 41

Figure 75: Model Full Assembly Bottom View ............................................................................... 41

Figure 76: Airfoil Inspection ........................................................................................................... 42

Figure 77: Model Reference Plane ................................................................................................. 43

Figure 78: Assembly Inspection 1 ................................................................................................... 43

Figure 79: Assembly Inspection 2 Figure 80: Assembly Inspection 3 ........................................ 43

Figure 81: Wiring Diagram (Side View) Figure 82: Wiring Diagram (Bottom View) .................. 44

Figure 83: Wiring Diagram (Overall) ............................................................................................... 44

Figure 84: Control System Block Diagram ...................................................................................... 45

Figure 85: Aileron Calibration ......................................................................................................... 45

Figure 86: Control Surface Loading Test Setup .............................................................................. 46

Figure 87: Linkage Schematic ......................................................................................................... 46

Figure 89: Aileron Load Test (1kg) Figure 90: Aileron Load Test (Close Up) ............................ 47

Figure 88: FBD Aileron .................................................................................................................... 47

Figure 91: Pre-Test Activities .......................................................................................................... 48

Figure 92: Model Balance Fitting Check ......................................................................................... 49

Figure 93: Roll Lock ......................................................................................................................... 49

Figure 95: Vertical Mounting Configuration Figure 96: Horizontal Mounting Configuration .... 50

Figure 94: Balance Calibration ........................................................................................................ 50

Figure 97: Wind Tunnel Test Block Diagram .................................................................................. 50

Figure 98: Mounting Configurations.............................................................................................. 52

Figure 99: Sign Conventions ........................................................................................................... 52

Figure 100: Velocity Sweep Test (10-64m/s) .................................................................................. 53

Figure 101: Velocity Sweep Data (10-40m/s) ................................................................................. 54

Figure 102: Pitch Step and Sweep Test .......................................................................................... 55

Figure 103: Yaw Step and Sweep Test ............................................................................................ 55

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III

Figure 104: Yaw Step and Sweep Test ............................................................................................ 56

Figure 105: Full Assembly (Isometric View).................................................................................... 96

Figure 106: Inspection Table 1 ....................................................................................................... 98

Figure 107: Inspection Table 2 ....................................................................................................... 98

Figure 108: Inspection Table 3 ....................................................................................................... 99

Figure 109: Inspection Table 4 ....................................................................................................... 99

Figure 110: Inspection Table 5 ..................................................................................................... 100

Figure 111: Inspection Table 6 ..................................................................................................... 100

Figure 112: Inspection Table 7 ..................................................................................................... 101

Figure 113: Inspection Table 8 ..................................................................................................... 101

Figure 114: SD7062 Airfoil Data (24) .............................................................................................. 89

Figure 115: 0010 Airfoil Data Generated From Airfoil Tool.com.................................................... 89

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IV

List of Tables Table 1: Permissible Measuring Error (3) ......................................................................................... 5

Table 2: NWTF at IITK (7) .................................................................................................................. 6

Table 3: Balance G Specification ....................................................................................................... 7

Table 4: Test Reynolds Number ........................................................................................................ 9

Table 5: Basic Dimension of Wind Tunnel Model ........................................................................... 11

Table 6: Horizontal Tail Stress Analysis .......................................................................................... 21

Table 7: Tail Mounting Block Fastener Analysis ............................................................................. 28

Table 8: Minimum Thread Engagement Length ............................................................................. 30

Table 9: Young's Modulus............................................................................................................... 30

Table 10: Factor of Safety Summary 1............................................................................................ 33

Table 11: Factor of Safety Summary (Joint Analysis) ..................................................................... 33

Table 12: SD7062 Max Ch ............................................................................................................... 37

Table 13: NACA0010 Max Ch (XFOIL) ............................................................................................. 38

Table 14: NACA0010 Max Ch (Analytical Approach) ...................................................................... 38

TTable 15: DS181 Specifications (13) ............................................................................................. 38

Table 16: Model Inspection Table .................................................................................................. 42

Table 17: Measuring Errors in Aerodynamic Coefficients .............................................................. 51

Table 18: VM and HM Data Comparison ........................................................................................ 51

Table 19: Material Properties Table ............................................................................................... 88

Table 20: Aerodynamic Force And Moment Derivatives................................................................ 90

Table 21: Standard Tightening Torque (15) .................................................................................... 92

Table 22: Hardware List .................................................................................................................. 96

Table 23: Model Mass Properties ................................................................................................. 103

Table 24: CG Table ........................................................................................................................ 103

Table 25: Material Density ........................................................................................................... 103

Table 26: Wind Tunnel Test Matrix ........................................................................................ 109

V

List of Symbols

Symbol Description Wing Area Wing span Vertical tail span / Angle of attack Cruise angle of attack Zero lift angle of attack Sideslip angle Elevator deflection Aileron deflection Rudder deflection Lift Coefficient at zero angle of attack Lift Coefficient with respect to angle of attack Lift Coefficient with respect to elevator deflection Incremental lift coefficient due to control surface deflection Pitching moment Pitching moment Coefficient Pitching moment Coefficient at zero lift Pitching moment Coefficient with respect to angle of attack Pitching moment Coefficient with respect to elevator deflection Yawing moment Yawing moment Coefficient Yawing moment Coefficient with respect to sideslip angle Yawing moment Coefficient with respect to rudder deflection Rolling moment Rolling moment Coefficient Rolling moment Coefficient with respect to sideslip angle Rolling moment Coefficient with respect to aileron deflection Side Force Side force Coefficient Side force Coefficient with respect to sideslip angle Side force Coefficient with respect to rudder deflection Hinge moment Hinge moment coefficient Hinge moment coefficient with respect to control surface deflection Bending Moment Torsion Tensile stress Ultimate tensile strength Yield tensile strength Shear stress Ultimate shear strength Yield shear strength Area moment of inertia Second moment of area Factor of safety

1

1. Introduction

1.1. Project Overview The Solar UAV project was initiated by the Defence Science Organization (DSO), National

Laboratories Singapore. This project aims at design and development of a medium altitude long

endurance (MALE) UAV with 24-hour continuous flight mission using solar energy as the sole

power source.

The solar UAV has a high wing and T-tail configuration. Low Reynolds laminar airfoil SD7062 was

selected as the wing airfoil while NACA 0010 was selected for the aerodynamic surfaces of the

empennage. The aircraft has a wing span of 17.7m and chord length of 1.25m which yield an

aspect ratio of 14.04. Having 85kg as the target weight, the UAV is designed to fly at 8000m

altitude with a target speed of 14m/s and the mission profile of the UAV is depicted in the figure

below. The latest estimated range and endurance of the UAV ;W IWWS Aero Loads, Performances, Stability and Control Analysi P; .

Figure 1: Mission Profile

The project consists of 5 main design teams namely:

1. Aerodynamics

2. Propulsion

3. Flight Controls and Avionics

4. Solar Energy

5. Aircraft Structures

In this detailed design phase of the solar UAV project, the Aerodynamics Design team will be

focusing on the study of the aerodynamics, flight stability and control, flight performance and

the propulsion system of the aircraft through three different aspects: empirical calculations,

computational fluid dynamics (CFD) and wind tunnel test. These data will be validated, compared

and analyzed to form the fundamental design principle for this solar UAV.

Subsequently, the Aerodynamics team is divided into 4 areas namely,

1. Aero Loads, Performances, Stability and Control Analysis: Part 1

2. Aero Loads, Performances, Stability and Control Analysis: Part 2

3. Wind Tunnel Testing

2

4. Computational Fluid Dynamics

1.2. Project Scope Particularly for wind tunnel testing, the objectives are to design and build a 1:9 scale wind tunnel

model of the solar UAV, to complete the servo selection for the actuation system of the model,

to do pre-test planning and wind tunnel test matrix preparation and lastly to perform wind

tunnel testing, aerodynamic data generation and preliminary analysis.

3

2. Literature Review

2.1. Model Design and Accuracy The type and construction of the wind tunnel model should be decided by the objective of the

tests and the tunnel facilities where the tests will be conducted. For example, 2D wing model is

used to study the aerodynamics of the 2D airfoil; half-model test is used for symmetrical object

and is normally focusing on the study of the particular aspect of the design such as engine and

wing flow interaction and acoustic condition near the wing. Half model test is also advantageous

in providing a more detailed representation than full models because larger scales can be used

(1). Two main important aspects to be considered in the model design are: 1. Safety

considerations; 2. model accuracy.

Figure 2: Half Model Test (1) Figure 3: 2D Wing Test

Laminated mahogany or other wood laminates are generally used for the construction of low

speed wind tunnel models. The laminated wood can withstand test speed up to 100mph without

metal reinforcement. For higher test speed, wood or various epoxy models with metal load

members are used. According to L-WWS WS TW TW, generally the criterion for model strength is structural deflection rather than yield load limits as high rigidity is desired. For

WWS S W W ;IIS NAA WS TW MSW W CW; ;aW a;I a a WWS H;WS ;W; ;W W ;S WW H;WS WS strength (2).

Besides, it is also advisable to have metal beams for the control surfaces in order to have best

precision in hinge line alignment. For wooden models, assembled components must be secured

using machined screws which normally mates with the threaded metal piece bonded into the

part. As the technology advances, new materials such as plastics, epoxy resins, fibreglass and

carbon fibre are also used as an alternative material for the construction.

An aircraft model with a span of 6 to 8 ft (1.83 to 2.44m) tested at 7 x 10ft tunnel requires a wing

contour accuracy up to 0.005in (1mm) to the actual contour; fuselage contour accuracy of

0.01in (2.5mm). Distinct ridges and joint on the model should be avoided. Furthermore,

surface finish of rms 10 (0.635micron) is desired for a metallic model (3).

4

Figure 4: Fibre-Glass Model (4) Figure 5: Laminated Mahogany Model NA-73X (5)

2.2. General Guidelines for Wind Tunnel Tests The guidelines and some of the standard procedure of a wind tunnel test are (reference taken

a L-WWS WS TW TW (3)),

1. Clearly address the problems and issues to be investigated and define the objectives of

the experiment. Clear objective statement is critical and essential in obtaining efficient

;I; a W W; aW; WSW ;S I ;SS W WW misunderstanding and confusion in any related information transferred.

2. Identify the required parameters to resolve the problem. For example, different

operating states or model configurations, geometries and etc.

3. Identify feasible model provisions and suitable test facilities. This will lead to the

preliminary design of the model and corresponding fixtures. From that, required wind

tunnel boundary corrections, tare, interference and other required data correction

needs to be identified.

4. Prepare the test schedules and runs (test matrix) including with configuration change

implications. Replication, randomization and blocking can be used to enrich the data

obtained.

5. Start the experiment and monitor all processes and data acquisition. This includes the

evaluation of the achieved accuracies and measurement precision.

6. Conduct data analysis and provide quantitative evaluation of the achieved data

accuracies. This information should be compiled and provided to the project personnel

such as aerodynamicists as a full data package.

2.2.1. General test procedure

Basic parameters of the tunnel facilities that needs to be taken care of before starting the tests

are: flow angularity, average dynamic pressure and balance loads. Flow angularity is the term

used to describe the flow angle with respect to the tunnel S; ; TW I method is by comparing the aerodynamic data of the normal and inverted mounted model.

Small amount of flow angularity is not evitable even in the most elegantly designed wind tunnels

5

(6). Therefore, flow angularity correction is required in the data analyses or even data acquisition

process to account for this issue.

The general procedure for the wind tunnel test activities are listed as follows,

1. First hand data check against predicted/estimated results should be done after the first

test run was completed. Some key parameters to be checked are 0, /, , 0 and 0 2. Determine the testing accuracy through repeatability tests. The repetition test can be

done immediately after one run or after several intervening runs. This is to determine

the reproducible accuracy of the settings including control surface deflections, accuracy

of the balance and speed control.

3. Plot out the acquired data promptly so that uncertain points can be substantiated and

make corrections for the following test runs. Bad data points and unexpected results can

also be spotted and adjustment and improvement on the system can be made

immediately.

4. Repeat a basic run periodically to identify any possible model warp or any time

dependent changes. It is also known that long hours of test would have certain effect on

the surface of the model. This check could help to identify this effect.

5. Create self-contained data sheets. Each data sheet should include the model designation,

configuration, test speed, date, tunnel temperature and pressure. It is also advisable to

have information like test Reynolds number and model dimensions stated in the data

sheet.

6. Have an accurate log of everything that happens during the test phase.

7. Keep the test runs list in a chronological order with clear indications of orders. Avoid

ambiguous or complex ordering system.

2.2.2. Permissible Measuring Errors of the Load Balance

AIIS W L WWS WS TW TW W WHW W; W a W ;S balance are stated as follows,

Low Angle of Attack High Angle of Attack

Lift = 0.001 or 0.1% = 0.002 or 0.25% Drag = 0.0001 or 0.1% = 0.002 or 0.25% Side Force = 0.001 or 0.1% = 0.002 or 0.25% Pitching Moment = 0.001 or 0.1% = 0.002 or 0.25% Yawing Moment = 0.0001 or 0.1% = 0.001 or 0.25% Rolling Moment = 0.001 or 0.1% = 0.002 or 0.25% Table 1: Permissible Measuring Error (3)

This table is used as a guideline only as the desired accuracies of the measurement differ

according to the test objectives and different test setups.

6

3. National Wind Tunnel Facilities The wind tunnel test was conducted at the National Wind Tunnel Facilities (NWTF) in Indian

Institute of Technology Kanpur (IITK), India. The following table provides a brief summary of the

facility:

Test Section 2.25m(H) x 3m(W) x 8.75m (L)

Fan Power Single, 1000kW

Highest Test Speed 80m/s

Reynolds Number Range (per m) 0 - 500000

Free-stream turbulence level U U Data Acquisition Systems PXI system with real time embedded controller

Virtual Instrumentation (LabView) Separation of console and user stations Standard DAQ and motion control hardware

Capabilities Full model testing with sting support system Half model testing with external balance Turntable system Moving belt for ground effect simulation Gust and cross wind simulation Aero-acoustic testing Laser light sheet generation system for flow visualization 3D stereoscopic PIV system

Table 2: NWTF at IITK (7)

LabVIEW is used for all the tunnel controls and data acquisition functions. It is a graphical

programming platform for measurement and control system developed by National Instruments

(8). The figure below shows the schematic of the communication between the DAQ system and

the instrumentation control interface. The DAQ system has a sampling rate of 200/sec.

Figure 6: DAQ System

In this experiment, the design test speed is 63.4m/s and turn table rig was used. More

information about the balance selected will be discussed in the next two sections.

7

3.1. Balance Selection and Balance Mounting Scheme There are seven different internal balances available for the wind tunnel facilities. The balances

are different in size and load range, targeted for different experimental purposes. The list of the

NWTF balances available is shown in the Appendix A.

Balance Selection was completed in the last year project. Based on the maximum lift coefficient

of the airfoil = 1.2 and dynamic head of 2205 at 60m/s, maximum load on the model was estimated as 695N. As a result Balance G was selected. The specification of the balance is

shown as follows,

Balance G

Dimension 30mm (diameter), 270mm (length)

Normal Force (N1,N2) 75kg (736N)

Side Force (S1, S2) 50kg (491N)

Axial Force (Ax) 20kg (196N)

Rolling moment (Rm) 4kgm (39Nm) Table 3: Balance G Specification

Figure 7: Balance G

The interface between the force balance and the model is very critical. Tight fitting must be

ensured at the interface attachment of the two components so that load transfer can be

achieved effectively and to prevent vibration or any mechanical interference in the assembly.

The figure below shows the balance G assembly with the interface units. The yellow component

in the front is the balance front interface unit which will be tightly fitted into the fuselage cap. It

is the sole contact point and hence the load transfer path from the model to the balance. 5mm

clearance was provided between the balance and model. The rear interface unit provides the

connection with the model strut and the entire balance assembly is mounted on the sting of

8

turntable shown in blue. The portion of the strut highlighted in red is the wiring guiding slot for

the wires of balance and the model actuation system.

Figure 8: Balance G Assembly Figure 9: Balance Interface Unit

Figure 10: Balance G Setup

The fuselage was made longer in order to accommodate the length of the balance assembly.

13mm (equivalent to 0.12m increase in the prototype) was added ahead of the wing leading

edge while 138mm was added behind the leading edge (equivalent to 1.24m increase in the

prototype). This geometric distortion would cause marginally higher drag in the model than the

actual prototype. The drag experienced by the actual size fuselage can be accounted for in the

post data analysis.

9

4. 1:9 Scale Wind Tunnel Model Design According to similarity law, similar flow pattern can be predicted in different flow conditions if

same Reynolds number is attained in the flow. Therefore, in wind tunnel practices, to have a

good resemblance of the actual flight condition, it is always desirable to have the test Reynolds

number that is similar to flight Reynolds number. However, this is often not the case due to the

limitations of the setup and the interaction between the model dimension and the tunnel

capabilities.

UAV Prototype 1:9 scale Wind Tunnel Model

Air Density (kgm-3) 0.5258

1.225

Dynamic Viscosity (Pas) 1.527x10-5

1.789x10-5

Speed (m/s) 14

63.4

Wing Chord (m) 1.26 0.14

Reynolds Number 607407 607775 Table 4: Test Reynolds Number

In this case, the model was designed to be tested at wind speed of 63.4m/s, which is achievable

in IITK wind tunnel facility, in order to attain the similar flow condition as the actual flight.

4.1. Design Evolution In the last year project, a conceptual wind tunnel model was created. Using aluminium as the

main material, the metallic model was designed to be modular for the ease of transportation and

assembly of the model. The model also possessed multiple interchangeable dihedral wing

brackets which provide different dihedral configurations for the study of the dihedral effect on

the flight stability of the aircraft.

Figure 11: Conceptual Model

10

Figure 12: Final Design

This year, the design of the model had gone through several major changes and improvements.

First of all, the material of the wing and horizontal tail were changed into carbon fibre and foam

core sandwich structure. CFRP and foam core structure is a better solution as compared to CNC

aluminium because of its lower cost of production and higher strength to weight ratio. On the

other hand, different dihedral configuration design on the CFRP wing was much more complex

and challenging. As the result, the idea of having multiple dihedral brackets had to be dropped.

Moreover, several joints in the model were improved and made feasible for the manufacturing

process such as, the joint between the tail boom and the empennage, fuselage and wing, vertical

fin, horizontal tail and etc. Several sharp corners in the previous design were also made

streamline in the current model.

Lastly, the dimensions of the aircraft had also been through several iterations. For example,

increase of the wing chord from 1.15m to 1.26m, fuselage diameter from 0.40 to 0.45m and etc

(refer to table for basic scaled down dimension of the wind tunnel model). Every change in

dimension or geometry of the aircraft requires revisit to the wind tunnel model design in terms

of manufacturing feasibility, structural integrity and cost consideration. After a long iterative

design process, the model was finally completed and was sent to IITK for wind tunnel test at the

end of March 2014.

Wing (SD7062) Solar UAV Wind Tunnel Model

Aspect ratio 14.04 14.04

Span 17.70 1.96667

Chord 1.26 0.14000

Wing area 22.30 0.27533

Aileron span (per side) 2.21 0.24583

Aileron chord (total) 0.25 0.02800

Horizontal Tail (HT) NACA 0010

Aspect ratio 5.00 5.00000

Taper ratio 1.00 1.00000

Span 3.34 0.37103

11

Chord 0.67 0.07421

Elevator chord (total) 0.20 0.02226

Elevator span (total) 3.34 0.37103

Vertical Fin (VF) NACA 0010

Aspect ratio 1.50 1.50000

Taper ratio 0.49 0.48700

Span 1.53 0.17003

Tip chord 0.67 0.07421

Root chord 1.37 0.15238

Rudder span (total) 1.30 0.14444

Rudder chord 0.41 0.04571

Fuselage

Infront of wing 0.80 0.08889

Behind of wing 0.40 0.04444

Total length 2.46 0.27333

Diameter 0.45 0.05000

Boom

1/4 chord of wing to 1/4 chord of horizontal tail 6.00 0.66667 Table 5: Basic Dimension of Wind Tunnel Model

4.2. Model Component Description In this section, the description of the design of each model component is presented. The detailed

3-view drawings of the design are attached in Appendix B.

Fuselage Cap (Aluminium)

Figure 13: Fuselage Cap Figure 14: Fuselage Cap (3-view Drawing Extracted)

Figure 14 shows sectional view of fuselage cap. The joint between the cap and the fuselage is

achieved by the M40 fine male thread on the fuselage cap and female thread at the inner wall of

the fuselage. As shown in the figure, the cylindrical cavity with a diameter of 35mm is designed

to accommodate the front interface unit (FIU) of the force balance. There is an internal tapered

slot matching the tapered tip of balance G at the FIU to ensure tight fit between these two

components. An M10 screw is used to secure the model to the force balance from the tip of the

cap. Another design features on the fuselage cap are the roll locks which are located radially

around the fuselage cap. Four M3 set screws (roll locks) were used to exert normal force onto

the FIU in order to prevent any rolling motion of model.

12

Fuselage (Aluminium)

Figure 15: Fuselage Figure 16: Fuselage (3-view Drawing Extracted)

The figures above shows the isometric and section view of the fuselage. The hollow section

provides 5mm clearance between fuselage inner wall and the force balance in order to avoid any

fowling between the two components. A 29mm x 130mm slot at rear of the fuselage permits to

take model strut from the aft interface unit (AIU) of the balance. The rectangular platform on the

fuselage, as shown in Figure 15, is the provision for the wing mounting block. The wing mounting

block sits on the fuselage and is secured by four M5 screws. Moreover, 3 etched marks (model

reference points) are made at top and both sides of the fuselage.

Wing Mounting Block (Aluminium)

Figure 17: Wing Mounting Block Figure 18: Wing Mounting Block (3-view Drawing Extracted)

Wing mounting block is one of the main structural components of the model. This component is

secured onto the fuselage by four M5 screws from the top. The wing will be sitting on the 2

degree angled surface of the mounting block and another six M5 screws are used to secure the

wing onto this component.

Wing (CFRP And Foam Core Sandwich Structure)

Figure 19: CFRP Foam Core Composite Wing

13

Figure 20: Wing Bottom View

The wing of the model is made out of carbon fibre and foam core sandwich structure. Two PVC

control hinges and three metal plates (two servo plate and a wing central plate) with threaded

holes are embedded in the wing in the carbon fibre laying up process. Figure 20 shows the

bottom view of the wing in which the wing central plate is made visible (shown in grey).

On the other hand, servo footprint cut out slots are made at both sides of the wing so that the

servos are made flush with the bottom surface of the wing and flow obstruction can be

minimized.

In addition to hard foam core, CFC skin and embedded metallic plate, each side of the wing is

further reinforced by 3 carbon flats extending from the central metallic plate to the tips.

Structural strength and stiffness due to such a spanwise reinforcement by carbon flats is not

quantified and taken into consideration in stress analysis and wing tip deflection studies.

Tail Boom Connector

Figure 21: Tailboom Connector Figure 22: Tailboom Connector (3-view Drawing Extracted)

This component serves as a connection between fuselage and tail boom. Tail boom is inserted

through the central 13mm-diameter circular slot as shown in Figure 22. A rectangular cut-out is

created for the unique mate with the flat tongue created at one end of the tailboom.

14

Fuselage Patch (Stainless Steel)

Figure 23: Fuselage Patch Figure 24: Fuselage Patch (3-view Drawing Extracted)

This component serves as a special nut plate for the M4 screw which secures the fuselage, tail

boom connector and the tail boom in place.

Tail Boom (Stainless Steel)

Figure 25: Tailboom

As discussed before, a flat tongue is created at one end of the tail boom in order to mate with

the horizontal surface at the tail boom connector. This prevents any undesired tilt of the

empennage which is mounted perpendicular to the tail boom. In addition, the tail boom is

designed with transitional fit (K7/h6 ISO standard) with the corresponding slots at the tail boom

connector and the empennage III. This transitional fit is a compromise between clearance and

interference fit. As frequent assembling and disassembling of the empennage is expected, the

transitional fit provides tight and firm fitting at the same time ensures smooth installation.

Empennage III (Aluminium)

Figure 26: Empennage III Figure 27: Empennage III (3-view Drawing Extracted)

Empennage III is the component onto which the vertical and horizontal tail assembly is fastened.

The front and aft fastener holes are the provisions for two M4 screws which are used to secure

VT Empennage Loft, a component on which the vertical fin will be attached, onto the

empennage III. The rear end of the tail boom is inserted through the 13mm diameter circular slot

in this component. Similarly, 3 etched marks are made on both sides of the component.

15

VT Empennage Loft (Aluminium)

Figure 28: VT Empennage Loft Figure 29: VT Empennage Loft (3-view Drawing Extracted)

This component serves as a special base plate for the vertical fin. On top of it, an airfoil footprint

is created to mate perfectly with the fin. The cross section gradually morphs into the footprint of

the empennage III at the bottom of the base plate. The two holes at the two ends are the

provisions for the M4 screws which are used to secure this component onto empennage III. The

other two counterbore holes in the middle are designed for the fasteners securing the fin.

Vertical Fin (Aluminium)

Figure 30: Vertical Fin

Similar to wing, there are two servo footprint cut out slots at the fin. One for the rudder another

for the elevator control. As the elevator servo is located at the fin, a special rocker arm linkage is

required to translate the servo horn motion in the yz-plane to elevator control horn motion in

the xy-plane. The two M2 tapped holes near the fin tip is designed for the installation of this

special linkage.

16

Tail Mounting Block (Aluminium)

Figure 31: Tail Mounting Block Figure 32: Tail Mounting Block (Side View)

The tail mounting block is mounted on the tip of the vertical fin by two M4 fasteners at the

central column. Four 3mm through holes are the provisions for the fasteners which secure the

tail to the mounting block. Four corners of the mounting block are filleted to reduce form drag.

The slot at the leading edge of this component provides reference for alignment during the

assembly process. Moreover, as shown in Figure 32, thin layer of material is removed at the

trailing edge to prevent interference with the horizontal tail.

Horizontal Tail (CFRP and Foam Core Sandwich Structure)

Figure 33: Horizontal Tail

Figure 34: Horizontal Tail Bottom View

Akin to wing, the tail is made out of CFRP and foam core sandwich structure with metal central

plate and PVC control hinges embedded. The tail is sitting on the tail mounting block which is

secured firmly at the tip of the vertical fin.

Rocker Arm

Figure 35: Rocker Arm Assembly 1 Figure 36: Rocker Arm Assembly 2

17

As discussed before, the rocker arm is a special linkage that translates the servo arm motion in the yz plane into WW; I I W ;W I I a components, rocker arm and nut plate. The aluminium nut plate is first mounted onto the fin

and then the rocker arm is mounted on top of that. Push rods from the servo and control horn

are then connected to the rocker arm to complete the assembly.

3D Printed Components (Polycarbonate)

The 3D printing technique has helped to ease the manufacturing process of the components

with more complex shape and smaller size. Although the achievable surface finish of the 3D

printed material is incomparable with both the well polished metallic and glossy CFRP surface, its

streamlined shape which can be easily formed using this advanced manufacturing technique is

significantly effective in reducing form drag caused by straight and sharp corners.

Wing and Horizontal Tail Cap

Figure 37: Wing Cap Figure 38: Horizontal Tail Cap

The wing and tail caps are pasted at the tips of their respective aerodynamic surfaces.

Control surfaces

Figure 39: Elevator

All the control surfaces of the model are 3D printed. Using this rapid-prototyping technology, the

slots for the control hinges can be easily made, as shown in the figure above. With these nicely

made slots, the hinges can be firmly inserted.

Levelling Block and Model Stand Design

Levelling blocks are used for the measurement and alignment purposes during the model

assembly and inspection stage. For example, the wing levelling block has an airfoil curvature that

mates the top surface of the wing. On top of the block, horizontal surface was created for the

placement of the inclinometer for angle measurement. The small rectangular embossed surface

W HI WW ; ; SW a W WW W W S W ; ;W IWI (see Figure 80).

18

Figure 40: Levelling Block and Model Stand Figure 41: Model On The Stand

The model stand is made up by three components: base, front stand and rear stand. The V-shape

at the front stand supports by the outer cylindrical surface of the fuselage while the rear stand

slots into the rectangular cut-out at the rear-end of the fuselage to prevent the model from

turning.

19

5. Model Stress Analysis

5.1. Mechanical Properties of Materials and Aerodynamics Inputs

on Component Loads

Appendix C includes mechanical properties of all the materials and the aerodynamic force and

moment derivatives used for estimating model loads.

5.2. Factor of Safety According to NAA WS TW MSW W CW; (2), W a;WW a;I a ;aW be equal or greater than 4 when using the material ;W WW W W WS strength is used. Furthermore, the thread engagement length should be at least one times the

nominal diameter of the fastener for tapped holes in material with ultimate tensile strength

greater than 830MPa. For tapped holes in materials less than 830MPa ultimate tensile strength,

1.5 times the nominal diameter of the fastener should be used as the minimum thread

engagement length. If less thread engagement is used, the minimum shear strength of the

threads in the joint shall be at least 4/3 times the bolt preload.

5.3. Stress Analysis of Model Structural Components In this section, the critical loading cases of the major load bearing structural components of the

model are evaluated.

5.3.1. Wing

The critical loading case of the wing occurs when = and with aileron fully deflected ( = 10) which yields maximum rolling moment on the aircraft. Rolling moment due to side slip is negligible in this calculation as the angle of side slip derivatives on rolling moment is very

insignificant as compared to aileron control derivatives on rolling moment.

Lift on wing is assumed to be elliptically distributed and the lift is acting at the centroid of the

quarter-WW H) on either sides of the wing. With a moment arm of 0.417m, the maximum bending moment due to lift occurs at wing root. The additional bending moment due

to rolling moment should also be taken into consideration to obtain the maximum bending

moment at the wing root.

Figure 42: Wing Bending Moment

The moment arm is calculated as follows,

0.417m

L/2

20

= 43 = 0.417

Where is half of the wing span = 0.984m Wing lift, = = 728 Hence, bending moment due to wing lift is =

2 = 152

Rolling moment coefficient of the model can be obtained as follows, = + = 0.028 Rolling moment due to roll and yaw rate are neglected here as we are only interested in static

motion. As discussed above, the rolling moment due to sideslip is also negligible.

Rolling moment can be obtained as follows, = = 33 And the resultant bending moment on the wing will be the sum of these two components

Conservatively, it is assumed that the bending moment is only sustained by the wing skin.

Assuming wing as cantilever beam structure, the formula of flexural stress on the wing due to

bending moment is shown below, = Unlike the regular cross-section of the cantilever beam, the area moment of inertia I of an airfoil

is relatively complex to obtain. Hence, in this analysis, the cambered airfoil is approximated as a

symmetrical airfoil with the same maximum thickness as shown in the figure below.

Figure 43: Airfoil Approximation 1

The area moment of inertia of the wing skin is computed in SolidWorks, = 6.26 1094; = 0.141402

= 9.8 The resultant flexural stress experienced on the wing, = 290 Factor of safety, = = 5.17 > 4

x

0.1c 0.8c

y

21

Direct shear due to transverse load and torsional shear due to wing pitching moment constitute

the total shear stress of the wing.

Bending shear stress,

= 2 = 0.22 Where = 1.62 1032 Pitching moment coefficient of the wing can be calculated as follows, = 0 + = 0.1 0.00570 10 = 0.157 Hence the resultant torsion due to pitching moment on the wing is, = = = 13.34 Shear stress borne by the wing skin is calculated using Bredt-Batho formulation, as shown below, = = 2 Where is the cross sectional area of the wing (1.68 1032); is the thickness of the 2-ply carbon fibre skin(0.8 103) = 4.96 = 5.14 > 4 Shear strength of the Carbon Fibre skin is 25.5MPa, experimental data by the structures team.

5.3.2. Horizontal Tail

The loading on the horizontal tail is very similar to that of the wing. By using the same approach,

the result of the analysis can be concluded in Table 3,

HT skin

Tensile stress (Bending) (Mpa) S.F

5.49 273

Shear Stress (Pitching moment) (MPa)

1.59 16

Table 6: Horizontal Tail Stress Analysis

5.3.3. Tail boom

Tail boom is also a critical structural component of the model. As depicted in Figure 44, tailboom

experiences torsion, sideward (lateral) and upward (longitudinal) bending moment due to

aerodynamic loads from the horizontal tail and vertical fin. As the result, the total stress

experienced at the tailboom due to the overlapping of all these stresses can be very significant.

The critical loading case (bending) for the tail boom occurs at the directional test, when the

SW W ; a SSW SWaWI ~r

22

Figure 44: Tail Boom Stress Analysis

Lift coefficient at horizontal tail = = 0.0724 8.840 Hence, lift at tail, = = 39 N Bending moment due to horizontal tail, = = 14 Where, =Tail lift moment arm = 0.371. Tail pitching moment is zero due to symmetrical airfoil.

Side force coefficient at vertical tail

= + = 0.002687 12 + 0.00223 15 Hence side force at vertical fin, = = 40N Consequently, the side ward bending moment due to side force from the fin is, = = 14 Where, = Fin lift moment arm =0.356. Similarly, fin airfoil pitching moment is zero due to symmetrical airfoil.

23

Figure 45: Tail Boom Cross Section

(+)Plus sign indicates tension while (-) minus sign indicates compression. It can be observed that

there must be a quadrant which sustains the highest tensile stress, depending on the nature of

the sideslip and rudder deflection.

Total tensile stress on the tail boom, = 0 + 0 = 131 MPa Where = 1.40 1094; 0 = 6.5 103 Factor of Safety, = 8.7 > 4 TW II; ;S I;W II ; ;W; W W SW W ; th aileron fully deflected. The resultant torsion on the tail boom is the sum of the two moments (rolling

moment due to deflected aileron and torsional moment due to the fin side force).

Figure 46: Tail Boom Torsion Analysis.

Figure 46 ; W;W a W SW WS ; - ; W W SW positive aileron deflection. Hence it can be seen that the resultant torsion experienced on the

tail boom is the sum of these moments.

-

24

Assuming that the side forIW ;I W a a MAC W SW SW aIW shown as follows, = = 20 = = 1.5 Where S;IW a a MAC W S; ; a W H m From section 5.3.1., rolling moment due to aileron deflection is equals to 33 . Consequently, the total torsion of the boom = 34.5 and the resultant shear stress due to torsion is calculated as follows, = 0 = 80 Where = 2 = 2.8 1094 The total shear stress of the boom is calculated as follows,

= + 2 + 2 = 80.3 = 8.5 > 4 5.4. Analysis of the Structural Joints

5.4.1. Wing Joint

Figure 47: Wing Joint Illustration

Besides wing lift, wing root bending moment (along X-axis) and pitching moment at quarter

chord (along Y-axis) all cause tensile stress at the 6 M5 fasteners used to secure the wing

mounting block to the embedded wing central plate. Based on the loading direction, it can be

concluded that the two screws in the last row will be subjected to the highest stress.

Tensile load due to lift bending and rolling moment(1) and Tensile load due to lift(2) ,

M5 screws x6

Wing Central Plate

New Wing

Mounting Block

y

x

25

1 = 2 +0.045 3 = 1497.4 ; 2 = 6 = 121.4

Figure 48: Tensile Load Due To Bending And Rolling Moment Figure 49: Tensile Load Due To Lift

Tensile Load due to pitching moment,

Figure 50: Tensile Load Due To Pitching Moment

Figure 50 shows the free body diagram of the 3 M5 screws which secure the wing to the

mounting block. Assuming linear load distribution, each bolt load R1, R2 and R3 can be

calculated as follows,

211 + 222 + 233 = Where 1 = 0.005;2 = 0.02;3 = 0.035 ; 1 = 13 3; 2 = 23 3; = 13.34 (obtained from section 5.3.1. 3 = 142 Hence, the total tensile stress subjected by the M5 screws is = 1+2+35 = 89.65 Fastener Part Number: CBSTS5-8

Strength Class: SUS304 equivalent = 4.5 > 4

45mm

F1 F2

2xR3

L

Bending Moment

d3

d2 d1

M_wing 2xR2 2xR1

26

5.4.2. Wing Mounting Block Fasteners Analysis

As shown in Figure 51, four M5 fasteners are used to mount the wing mounting block on the

fuselage. The entire wing load and moment are concentrated here; the four mounting screws

sustain tensile load due to wing lift, pitching moment and rolling moment due to asymmetrical

lift (aileron deflection). Therefore it is critical to examine the structural integrity of this assembly.

Figure 51: Wing Mounting Block M5 Fasteners Figure 52: Tensile Load Due To Pitching Moment

Tensile load due to wing lift 1 = 182; Tensile load due to pitching moment is obtained using the same approach as in section 5.4.1. 2 = 597 = 0 + + = 0.1 0.00570 10 0.0223 15 = 0.492; = = 41.76; 2 = 597

Figure 53: Tensile Load Due To Rolling Moment

Taking the edge of the mounting block as the pivot point for the rolling moment, the tensile load

(3) borne by the central bolt can be found as follows, 3 = ( 0.025 )4 = 334 The maximum stress experienced on the central M5 screw and the safety factor can be found as,

M5 screws x4 L

Mwing

l

F2

Mwing

45mm

30mm

15mm

0.25c

F3

25mm

l

27

= 1+2+35 = 56.7 Fastener Part Number: CBSST5-8

Strength Class: A2-50 = 8.8 > 4 5.4.3. Horizontal Tail Joint Analysis

As discussed before, a central aluminium plate H; T; CW; P;W is embedded in the composite tail. As shown in Figure 54 the tail load will be transferred through the four M3

screws to the fin. The critical loading condition for the tail occurs at max AOA with full elevator

deflection (~e C;WS W WW W load on the tail is expected to be much smaller.

Note: HT Angle of attack - IWS ; C AOA T; W ;W -

Figure 54: Tail Joint Analysis Figure 55: HT Lift Bending = + = 0.0724 8.840 + 0.004835 15 = 0.713 = = 43 Using the similar analysis done for the wing joint, 1 = 11 and 1 = 43 Hence maximum tensile stress experienced on the M3 fasteners can be obtained as follows = 1+23 = 7.6 Fastener part number: CBSST3-5

Strength Class: A2-50 = 65.7 > 4

F1

M3 screws x4

Horizontal Tail

Central Plate LHT

F2

20mm

L/2

79mm

28

5.4.4. Tail Mounting Block Fasteners Analysis

Figure 56: Tail Mounting Block Fastener Analysis

Tail load is fully transferred to the two tail mounting block M4 fasteners. Using the similar

approach as the previous calculation, the results can be obtained as follows,

Fastener part number: CBSST4-8

Strength Class: A2-50

Tensile load (bending) (N) Tensile load (lift) (N) Total stress (Mpa) S.F.

0.00 21.5 1.71 292.2

Table 7: Tail Mounting Block Fastener Analysis

5.4.5. Vertical Fin Joint Analysis

The vertical fin is secured by two M4 fasteners at the root of the fin. The critical loading case for

W a;WW HW ; ; ;S W ~r A Figure 57, the fasteners sustain a combination of transverse shear and bending stress due to the offset side force from

the fin.

Figure 57: Fin Joint Analysis

M4 Fasteners x 2

75mm

Transverse force = 20; Tension due to tail lift 1 = 22 Bending moment = 0.075 = 1.5 Axial force due to bending moment

2 = 0.00752 = 192

Tensile stress on the bolt = 17

Fastener part number: CBSST4-12

Strength Class: A2-50

= 29.4 > 4

YVT

V

F1+F2

7.5mm

29

5.4.6. VT Empennage Loft Fastener Analysis

Figure 58: VT Empennage Loft Fastener Analysis

Fastener part number: KBBS4-10; FKBB4-26-6

Strength Class: A2-50; 8.8

TW W; ;S W W a W a H;W ;W VT EW;W La M a;WW IW ;W the same as the fin joint fasteners. The distance between the two screws is wider and therefore

it can be deduced that the resolved shear load will be smaller than the fin joint screws. Hence,

without showing detail calculations, it can be concluded that the results obtained from section

5.4.5. are enough for assuring the safety of the joint.

5.4.7. Tail Boom Joint Analysis

The tail boom of the model is only secured by one M4 fastener as shown in the figure below. The

bolt is subjected to transverse shear due to the combined torsion between the boom and the

fuselage.

Figure 59: Tail Boom Joint Analysis Figure 60: M4 Fastener Internal Load

The maximum torsion on the tail boom joint can be obtained from section 5.3.3. = 34.5 Hence transverse shear V and the corresponding shear stress can be calculated as shown, =

0.0127= 2717; = 216

Fastener part number: FBAB4-48-5

Strength class: 12.9

2 M4 Fasteners

V

V

30

= 3.4 > 3 (Yield Shear Strength is used) 5.5. Thread Engagement Length

Minimum Thread Engagement Length Le is calculated using the two formulas below,

For bolt shank to fail before thread stripping in the bolt (first order approximation) (9),

= 0.7454 0.938220.27125 0.54127

For bolt shank to fail before thread stripping in the nut (first order approximation) (9),

= 0.7454 0.938220.27125 0.54127

Where, =yield strength of the bolt material; = yield strength of the nut material The larger value of the above calculated Le should be used as the minimum thread engagement

length required.

Joint Le (mm) Thread engagement

length provided (mm)

Wing Mounting Block to Fuselage 3.45 4

Wing to Wing Mounting Block 3.45 5

Tail Boom to Fuselage 2.7 4.5

VT Empennage Loft to Empennage III 3.94 6

Vertical Fin to VT Empennage Loft 2.7 6

Tail Mounting Block to Vertical Fin 2.70 5

Horizontal Tail to Tail Mounting Block 2.05 3 Table 8: Minimum Thread Engagement Length

5.6. Wing and Tail deflection Wing tip and tail deflection due to aerodynamic forces are calculated here. Similarly,

conservative approximations are used in this analysis in order to simplify the mathematical

calculations as well as making sure that the actual deflection will be smaller than expected value.

Assumptions:

1. Wing and tail boom as approximated as cantilever beam (Bernoulli-Euler Beam

Model)

2. Rectangular wing lift distribution

3. Wing CFRP skin is the sole supporting structure of the wing

4. Aileron control effectiveness = 0.2 (aileron chord to wing chord ratio) 5. Wing area moment of inertia is approximated the same as in section 5.3.1.

Material Youngs Modulus E Carbon Fibre UD 135GPa AISI 302 Stainless Steel 193GPa

Table 9: Young's Modulus

31

TW ;HW ;HW SW W Y S a W WWIW ;W; a I;I; W deflection using Bernoulli-Euler Beam Model.

The critical wing tip deflection occurs at full aileron deflection at highest possible angle of attack.

However, deflecting aileron when = will lead to aileron stall. Therefore this analysis is WaWS ; AOA a SWaWIWS ;W ~a A W W W ;W W will have = where the rest of the part of the wing will have 1 = 0.093 7 = 0.651.

Figure 61: Wing Lift Distribution

R1 and R2 are the distributed load on the wing. 1 = 1 ; 2 = 1 = 12(0.98350.246) 20.9835 = 240/; 2 = 12(0.246) 20.9835 = 59/ The tip deflection can therefore be calculated as follows, = 1 { 132 0.2460 + 2 0.246 0.2462 + 10.24622 0.98350.246 }; = 25 The actual deflection is expected to be much smaller than the estimated value due to the

IW;W ;;I W a W; ISWWS ;S WI;; a distribution) used in this analysis. Furthermore, there are 6 carbon fibre strips as stiffeners (3 on

each side of the wing) will be bonded at the bottom wing skin to enhance the wing stiffness. This

unison of CFC skin, hard foam core, carbon rod stiffeners essentially constitute the structural

strength of the wing. Hence the actual deflection will be much lower than the values obtained

above.

0.246 0 0.9835

R1 N/m R2 N/m

32

Figure 62: Tail Boom Upward Bending Deflection

Similar approach is used in analyzing the tail deflection. = = 7.451; = 7; = 43 ( 5.4.3. );

Figure 63: FBD For Tail Deflection Analysis = 1 43 70.02015 + 2 7.45320.3710 = 2.4

Figure 64: Tail Boom Defection Due To Side Force

R N/m

Wemp

LHT

20.15mm

R N/m LHT-Wemp

20.15mm 371 0

YVT 0

371

33

Similarly, tail boom deflection due to side force is calculated as follows, = 1 2 0.3710 = 2.5 Where = 40 is the side force obtained from section 5.3.3.

5.7. Summary The tables below show a summary of the results from the model components and joint analysis.

It is concluded that al W II; I; IW ; W ; H a;I a ;aW are larger than 4 (for UTS) and 3 (for YTS).

Components Type of stress Factor of Safety

Wing Tensile 5.2

Shear 5.1

Horizontal Tail Tensile 273.0

Shear 16

Tail Boom Tensile 8.7

Shear 8.5 Table 10: Factor of Safety Summary 1

Joint Fastener size Type of stress Factor of Safety

Wing to Wing Mounting Block M5 Tensile 4.5

Wing Mounting Block to Fuselage M5 Tensile 8.8

Horizontal Tail to Tail Mounting

Block

M3 Tensile 65.7

Tail Mounting Block to Vertical Fin M4 Tensile 292.2

Vertical Fin to VT Empennage Loft M4 Tensile 29.4

Tail Boom to Connector M4 Shear 3.4 Table 11: Factor of Safety Summary (Joint Analysis)

34

6. Hinge Moment Estimation and Servo Selection Integrated model actuation system was used in this wind tunnel model so that control surface

deflection can be achieved remotely. This would significantly reduce the downtime of the

experiment and ensure a smoother run between each test. As the experiment would be

conducted at high speed 63m/s, much higher than most of the RC aircraft flight speed, the

selected servo must be strong enough to withstand the high pressure and actuate under the high

loading condition. Therefore servo selection becomes one of the critical design stages for the

wind tunnel model.

Some other selection criteria are as follows:

1. Slim, as the servos are required to be fitted in the respective aerodynamic surfaces,

protrusion of the parts into the flow is undesirable.

2. Fast response and precise actuation

3. Easy installation

In the servo selection process, hinge moments at the respective control surfaces were evaluated.

The servo models that meet the torque requirement were then being compared based on the

abovementioned selection criteria. Finally the suitable servo was selected.

Two different approaches were used to estimate the hinge moment of the respective control

surfaces. Suitable actuators were then be selected for the model control system based on the

calculated results. As the methods only give a rough gauge of the hinge moment magnitude,

conservative results were chosen as the baseline for the servo selection. This ensured that the

servo is slightly oversized and able to actuate efficiently in the wind tunnel environment.

6.1. XFOIL Computation: First developed by Mark Drela at Massachusetts Institute of Technology (MIT), XFOIL is a

program for the design and analysis of the low speed subsonic airfoils (10). In the stress analysis,

XFOIL was also used to calculate the pitching moment coefficient of the NACA0010 with

deflected control surfaces (elevator and rudder).

Figure 65: GDES Function

Flap at the desired hinge line of the 2D airfoil was modified in GDES, integrated airfoil design

; XFOIL U FMOM I;S W W W IWaaIW a W ;a

35

was obtained. However, two important keys to take note about this analysis were: 1.

Incremental hinge coefficient of the SD7062 is already available in experimental data, only hinge

moment due to angle of attack is computed. The total hinge moment coefficient is the sum of

the experimental data and XFOIL results. 2. The reference parameters used to non-

dimensionalize hinge moment in XFOIL are wing area and wing span. Extra calculation steps are

required to obtain the conventional hinge moment coefficient expression.

= 122

Where and are control surface area and chord respectively

Wing (SD 7062):

Hinge moment coefficient at different angle of attack (AoA sweep from - ;nd zero flap deflection angle) was estimated by XFOIL. The due to a; SWaWI ~ was obtained from the experimental data - EWW; ;S I;; S a a;WS ;a SW RWS HW TW a both dimensionless numbers yielded the total .

Figure 66: Experimental Data of() (11) Horizontal Tail and Vertical Fin (NACA 0010):

As no experimental data is available for NACA 0010, of the elevator and rudder were estimated purely by XFOIL. As stated before, flap position and deflection angle are set under

GDES function of XFOIL.

36

6.2. Empirical Approach Cl estimation This is a similar method to the XFOIL approach. However the hinge moment due to control

surface deflection was obtained empirically. Firstly, the incremental lift due to control surface deflection was estimated. It was assumed that this purely acted on the control surface and the centre of pressure is located at 50% of the control surface chord. The hinge moment was

then obtained by multiplying with half of the chord length of the control surface, which was the equivalent moment arm. due to deflected control surface =

Where, - flap effectiveness = 0.3 a 2D lift curve slope = 0.1096/degree

As discussed above, it was assumed that was completely acting on the control surface, with centre of pressure located at 0.5 of the control surface chord. Hence multiplying with 0.5 and dividing the term with control surface chord to non-dimensionalize it. = 0.5( 1)

Where, control surface chord

due to angle of attack The equation below shows the estimation of the incremental lift acting on the control surface at

different angle of attack. The coefficient 0.5 represents the fraction of the total lift increment

that acts on the control surface. = 0.5

Similarly, the hinge moment coefficient due to angle of attack was calculated as follows, = 0.5( 1)

Hence, the total hinge moment coefficient can be express as the following equation, = +

37

6.3. Results: This section presents a summary of the results obtained. Figure 67 and 68 show the relationship

between hinge moment coefficient and the control surface deflection angle using the XFOIL and empirical approach respectively.

Figure 67: Ch vs deflection angle (sigma)

Figure 68: Ch vs deflection angle (empirical)

It can be seen from both of the graphs that higher magnitude of moment occurs at higher angle

of attack and larger control surface deflection angle.

The following tables show the comparison between the results of SD7062 (maximum hinge

W IWaaIW a H;WS ; AOA ~ a SaaWW ;;IW XFOIL Empirical

Ch_max 0.96 0.319

HM_max (Nm) 0.40 0.150

HM_max (kgcm) 4.11 1.534

Table 12: SD7062 Max Ch

Conservatively, hinge moment of 4.11kgcm was chosen to be the wing servo selection criteria

(torque)

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 5 10 15

Ch

_to

tal

deflection angle_sigma (degree)

Ch vs sigma (XFOIL)

AoA -7deg

AoA -5deg

AoA -3deg

AoA 0deg

AoA 2deg

AoA 4deg

AoA 6deg

AoA 8deg

AoA 10deg

0

0.1

0.2

0.3

0 5 10 15

Ch

_to

tal

deflection angle_sigma (deg)

Ch vs sigma(Empirical)AoA -7

AoA -5

AoA-3

AoA0

AoA2

AoA4

AoA6

AoA8

AoA10

38

Similarly, the following tables show the comparison between the results of NACA 0010

(maximum hinge moment coefficient obtained at AOA ~ a SaaWW approaches.

NACA 0010 XFOIL Approach (Max Ch)

Rudder Ch_Max 0.0142

HM (Nm) 0.086

HM (kgcm) 0.877

Elevator Ch_Max 0.0142

HM (Nm) 0.0744

HM (kgcm) 0.758

Table 13: NACA0010 Max Ch (XFOIL)

NACA 0010 Empirical Approach (Max Ch)

Rudder Ch_Max 0.244

HM (Nm) 0.0179

HM (kgcm) 1.827

Elevator Ch_Max 0.244

HM (Nm) 0.111

HM (kgcm) 1.136

Table 14: NACA0010 Max Ch (Analytical Approach)

Conservatively, HM of 1.827kgcm was chosen to be the tail servo selection criteria (torque)

The selected servo is shown in the figure below.

Figure 69: JR DS181 Slim Wing Servo (Courtesy of

hobbyking.com (12))

T

Table 15: DS181 Specifications (13)

DS181 specifications

Type Digital

Torque 4.3kgcm@4.8V; 5.4kgcm@6V

Speed 16sec/60degree@4.8V;

13sec/60degree@6V

Size 30 x 10.9 x 29mm

Weight 20g

Gear

Type

Metal

Voltage 4.8V to 6.0V

mailto:5.4kgcm@6V

39

7. Manufacturing Process of the Wind Tunnel Model The wind tunnel model is made out of four major materials namely, CFRP and foam core

sandwich structure, 3D-printed polycarbonate, aluminium and stainless steel. The components

of the model were manufactured by different vendors who are specialized in the manufacturing

of the respective materials.

Most of the metallic components such as fuselage, tail boom and etc are created using milling

and lathing machine. Lathing machine was used to create the rounded surface at the fuselage

cap and tail boom connector. More complex shape like vertical fin are created using computer

numerical control (CNC) machine.

Figure 70: Metallic Components

(Left: Model body components; Right: Vertical Fin)

The figures below show the manufacturing process of the CFRP wing and horizontal tail. The

moulds were first cut out using the automated KUKA robotic arm. According to the vendor, the

machine has an accuracy of 0.5mm. Carbon fibre laying-up and vacuum bagging process were

then carried out at both of the male and female moulds. Metal plates, PVC control hinges and

carbon stiffeners were integrated in between the laying up process to ensure good attachment

with the skin. Foam core was sandwiched between the moulds and resin was pumped into the

mould to fill up the cavities. The wing and tail were finished by a thin layer of paint.

40

Figure 71: CFRP Wing and Tail

(Top left corner: Mould being cut out by KUKA machine; Top right: Horizontal Tail with central

plate integrated; Bottom: Finished wing)

Simple load test was carried out by the vendor to show the strength of the wing. 100kg load

(30kg more than the maximum aerodynamic load) was applied on the wing, around the effective

wing lifting point. The wing generally displayed high strength capability with smaller than

expected deflection at the wingtip.

Figure 72: Wing Load Test

41

8. Model Assembly and Inspection The figures above show the 3 view of the full assembly of the model. The complete assembly

steps and hardware list can be found in Appendix D.

Figure 73: Model Full Assembly Front View

Figure 74: Model Full Assembly Side View

Figure 75: Model Full Assembly Bottom View

Inspection of the individual component as well as the full model assembly is an important

process before the model enters the wind tunnel. This process is to identify the shape

conformity and any dimensional deviation of the physical model from the CAD design. Key

dimension such as wing span, chord, fuselage diameter and etc, of the actual model parts were

measured. In order to minimize human error in measurement, three separate measurements

were taken and averaged values of the three were used. The measured data is then compared

42

with the CAD data and the error percentage can be calculated. The following table shows a

portion of the result extracted from the model inspection stage table. Complete model

inspection table can be found in Appendix E.

CAD Measured

Wing Span(without wing caps) 1947.00 1947.50

Wing Chord 133.50 133.50

Aileron Span 246.00 246.00

Aileron Chord 27.80 28.00

Horizontal Tail Span (without tail caps) 351.00 352.00

Horizontal Tail Chord 51.95 52.00

Fuselage Diameter 50.00 50.00

Fuselage Length 310.00 310.00

Vertical Fin Root Chord 149.40 149.00

Vertical Fin Span 179.00 179.00 Table 16: Model Inspection Table

In general, the dimensional accuracy of the model fluctuates about 1mm. Nonetheless, most of

the key features such as wing span, chord and etc were found to be matching the CAD data

(refer to Table 16).

Besides, the shape conformity of the airfoil section of the wing and tail were checked by

sweeping an airfoil template acrylic block along the span wise of the respective aerodynamic

surfaces. Unlike the coordinate-measuring machine, error could not be quantified using this

method. However, it gives the user a gauge on how well the shape of the physical wing conforms

to the design. It can be seen from Figure 76, the cross section of the wing matches well with the

airfoil template, the gap between the two components is almost unnoticeable. This gives us

enough confidence that the wind tunnel model carries a high degree of resemblance to the

actual prototype.

Figure 76: Airfoil Inspection

As discussed in the previous section, model reference points were made on the model by

creating 0.5mm etched marks on the fuselage. Fuselage reference line and model reference

planes can be constructed for model assembly and inspection purposes by adopting the

reference points, as seen from the Figure 77. Similarly, the position of wing, tail and other X-

stations (relative position with respect to aircraft nose) were checked.