AN INVESTIGATION OF FLEXIBLE ROGALLO TYPE WING AERODYNAMICS AND THE APPLICATIONS TO UAVS
A Thesis
Presented in Partial Fulfillment of the Requirements for
the Degree of Master of Science in the
Graduate School of The Ohio State University
By
Matthew L. Warchol
*****
The Ohio State University 2009
Master’s Examination Committee: Approved By:
Dr. Gerald M. Gregorek, Advisor
Dr. Richard J. Freuler ______Advisor
Aeronautical and Astronautical Engineering Graduate Program
ABSTRACT
With the threat of biological warfare, the development of small unmanned aerial vehicles (UAVs) capable of flying missions for chemical assessment, also known as
“dirty” missions, has become a growing need. The practicality of flexible parawings with the advantages of reducing packaging size has been assessed in this work. Wind tunnel testing of a highly flexible delta wing with a leading edge sweep of 30 was performed to see the effects on aerodynamic performance as a function of decreasing wing billow. A ratio of one half of the wing wetted area to wing planform area was used to define the amount of billow between wind tunnel runs. The results gave the highest value of
CL max 0.6 on the wing configuration with a billow parameter of 1.
Based on initial tests, additional investigation on increasing the structural robustness of the flexible wing was performed to improve the aerodynamic performance.
This was done by adding wing ribs between the wing leading edge and fuselage. Results showed that increasing the robustness would significantly improve the aerodynamic
performance. A CL max 1.2 was achieved during the wind tunnel analysis at configurations with 3 and 7 ribs per wing.
A full wind tunnel analysis was done on a model with an integration of the improve wing configuration, fuselage shape, and canard design. The testing of the UAV system included analysis of the aerodynamic performance as well as the longitudinal and ii lateral-directional stability and controllability with variations in leading edge sweep from
15 to 60 by 15 increments. A CL max 1.51 was achieved with the model
configuration having 15 and decreased to a CL max 1.1 with 45 .
The UAV’s performance coefficients were found at cruise conditions for each model configuration from the test results and used with the aerodynamic analysis to predict the respective UAV flight characteristics. The performance predictions included topics on thrust horsepower, range, endurance, aircraft structure, turning flight, climbing flight, and gliding flight. Three potential mission scenarios are described in detail with full mission profiles to complete each scenario.
iii
Dedicated to my dog Havoc
iv
ACKNOWLEDGMENTS
I express my sincerest gratitude for my advisor Dr. Gerald M. Gregorek, for the knowledge that he has shared, the opportunities that he has provided, for the guidance he has given, and encouragement in the practiced engineering society. I want to thank Cliff Whitfield for all the assistance and direction he was provided during the completion of this project. Thanks to Mr. Jerome Pearson and Star Technology and Research for the opportunity to contribute, expand personally and professionally with in the engineering community. I give my deepest respect and gratitude for Dr. Rick J. Freuler and Dr. John D. Lee, for the experimental and professional experiences that they have provided. Thanks to Dr. Jolanta M. Janiszewska for her assistance and guidance. And, most admiration for my family: without their support and encouragement, this would not have been possible. Also, to the unacknowledged that have in some way assisted or guided me thus far, I thank you.
v
VITA
October 25, 1983…………………. Born, Toledo Ohio, USA
June 10, 2007.……………………. B.S., Aeronautical and Astronautical Engineering, The Ohio State University
September 2007 – Present………… Graduate Research/ Teaching Assistant, Aeronautical and Astronautical Engineering, The Ohio State University
FIELDS OF STUDY
Major Field: Aeronautical and Astronautical Engineering
vi
TABLE OF CONTENTS
ABSTRACT……………………………………………………………... ii DEDICATION………………………………………………………….. iv ACKNOWLEDGMENTS……………………………………………… v VITA…………………………………………………………………...... vi LIST OF EQUATIONS………………………………………………… x LIST OF FIGURES…………………………………………………….. xi LIST OF TABLES……………………………………………………… xiv NOMENCLATURE…………………………………………………….. xvi
CHAPTERS:
1. INTRODUCTION ……………………………………………… 1
2. EXPERIMENTAL DESIGN / DEVELOPMENT / SETUP…. 4 2.1 Design Considerations……………………………………….. 4 2.2 Experimental Facility ………………………………………. 5 2.3 Wind Tunnel Model…………………………………………. 6 2.3.1 Series 1 Wind Tunnel Model………………………. 6 2.3.2 Series 2 Wind Tunnel Model ……………………… 7 2.4 Instrumentation………………………………………………. 9 2.4.1 Hardware…………………………………………… 9 2.4.2 Data Acquisition and Reduction…………………… 13 2.5 Test Procedure……………………………………………….. 14
vii 3. SERIES 1 - FLEXIBLE WING INVESTIGATION AND RESULTS 16 3.1 Wind Tunnel Investigation…………………………………... 18 3.1.1 Effects of Billow…………………………………… 18 3.1.2 Effects of Reducing Wing Deformation ………..…. 20
4. THE UAV TODAY…………………………………………….. 24
5. SERIES 2 - FLEXIBLE WING AND FUSELAGE INVESTIGATION AND RESULTS…………………………………………………. 26 5.1 Wind Tunnel Investigation…………………………………... 28 5.1.1 Effects of Leading Edge Sweep…………………… 28 5.1.2 Effects of Canard Deflection………………………. 32 5.1.3 Effects Due to Sideslip……………………………... 38 5.2 UAV Configuration………………………………………….. 40 5.2.1 15 Leading Edge Sweep…………………………... 41 5.2.2 30 Leading Edge Sweep………………………….. 42 5.2.3 45 Leading Edge Sweep………………………….. 43 5.2.4 60 Leading Edge Sweep………………………….. 44 6. UAV PERFORMANCE ANALYSIS……………………….... 45 6.1 Performance………………………………………………….. 47 6.1.1 Thrust Horsepower Required………………………. 47 6.1.2 Thrust Horsepower Available……………………… 48 6.1.3 Range and Endurance……………………………… 49 6.1.4 Rate of Climb………………………………………. 51 6.1.5 Gliding Flight……………………………………… 52 6.1.6 V-n Diagrams………………………………………. 53 6.1.7 Level Turns ……………………………………….. 57 6.2 Mission Scenario and Profiles……………………………….. 59 6.2.1 Mission 1…………………………………………… 59 6.2.2 Mission 2…………………………………………… 60 viii 6.2.3 Mission 3…………………………………………… 62 7. SUMMARY AND CONCLUSION……………………………. 65
REFERENCES………………………………………………………….. 67 APPENDIX A……………………………………………………………. 68 APPENDIX B……………………………………………………………. 71
ix
LIST OF EQUATIONS
3.1 Billow Parameter…………………………………………………. 17
5.1 Total Drag………………………………………………………… 30
5.2 Mean Aerodynamic Chord………………………………………. 40
5.3 Taper Ratio………………………………………………………. 40
6.1 Thrust Horsepower Required…………………………………….. 47
6.2 Thrust Horsepower Available……………………………………. 48
6.3 Rate of Climb…………………………………………………….. 51
6.4 Minimum Glide Angle…………………………………………… 53
6.5 Maximum Glide Range………………………………………….. 53
6.6 Maximum Load Factor…………………………………………… 54
6.7 Gust Loads……………………………………………………….. 54
6.8 Alleviation Factor………………………………………………… 55
6.9 Mass Ratio………………………………………………………... 55
6.10 Turn Radius………………………………………………………. 57
6.11 Turning Rate……………………………………………………… 57
6.12 Minimum Turn Radius…………………………………………… 58
6.13 Maximum Turning Rate………………………………………….. 58
x
LIST OF FIGURES
Figure 2.1: AARL 3’ x 5’ Eiffel Type Subsonic Wind Tunnel with Model Installed 5
Figure 2.2: Series 1 Wind Tunnel Model - 30 ……………………. 6
Figure 2.3: Series 2 Wind Tunnel Model - …………………… 7
Figure 2.4: Six Component Balance…………………………………… 9
Figure 2.5: Power supply and Amplifiers……………………………… 10
Figure 2.6: Normal Force Calibration .………………………………… 11
Figure 2.7: Stability Axis Force Components………………………….. 12
Figure 2.8: LabVIEW Virtual Instrument Front Panel…………………. 13
Figure 3.1: Wing model………………………………………………… 17
Figure 3.2: Lift coefficient vs. Angle of attack………………………… 18
Figure 3.3: L/D vs. Lift Coefficient with Wing Deformation …………. 19
Figure 3.4: Drag Polar with Wing Deformation……………………….. 20
Figure 3.5: Wind Tunnel Model with 1 Rib with Wing Deformation….. 21
Figure 3.6: Lift Coefficient vs. Angle of Attack with Rib Addition …… 21
Figure 3.7: Drag Polar with Rib Addition……………………………… 22
Figure 3.8: L/D vs. Lift Coefficient with Rib Addition ………………… 23
Figure 5.1: Wing and Fuselage model…………………………………… 27
Figure 5.2: Lift coefficient vs. Angle of attack with Varied Leading Edge Sweep 28
Figure 5.3: Drag Polar with Varied Leading Edge Sweep ……………. 29
xi 2 Figure 5.4: C L vs. CD with Varied Leading Edge Sweep ……………… 30
Figure 5.5: L/D vs. Lift Coefficient with Varied Leading Edge Sweep … 31
Figure 5.6: Pitching Moment coefficient vs. angle of attack……………. 32
Figure 5.7: Pitching Moment coefficient vs. angle of attack…………….. 33 ( 15 , for elevator deflection)
Figure 5.8 Pitching Moment coefficient vs. angle of attack……………... 34 ( 30 , for elevator deflection)
Figure 5.9: Pitching Moment Coefficient vs. Angle of Attack………….. 35
( 60 , for elevator deflection)
Figure 5.10: Pitching Moment coefficient vs. angle of attack………….. 35 ( 45 , for elevator deflection)
Figure 5.11: Rolling Moment for Aileron………………………………. 36
Figure 5.12: Side Force vs. Aileron Deflection………………………….. 37
Figure 5.13: Yawing Moment for Aileron Deflection…………………… 37
Figure 5.14 Yawing Moment vs. Sideslip………………………………. 38
Figure 5.15 Rolling Moment vs. Sideslip……………………………….. 38
Figure 5.16 Side Force vs. Sideslip ……………………………………. 39
Figure 5.17 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle
Geometry -15 Leading Edge Sweep……………...... 41
Figure 5.18 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle
Geometry -30 Leading Edge Sweep……………………………. 42
Figure 5.19 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle
Geometry - 45 Leading Edge Sweep……………………………. 43
xii Figure 5.20 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle
Geometry - 60 Leading Edge Sweep……………...... 44
Figure 6.1: Thrust Horsepower Required………………………………… 47
Figure 6.2: Sea Level Rate of Climb……………………………………… 52
Figure 6.3: V-n Diagram with Gust Lines for 15 ……………………. 55
Figure 6.4: V-n Diagram with Gust Lines for 30 ……………………. 56
Figure 6.5: V-n Diagram with Gust Lines for 45 ……………………. 56
Figure 6.6: V-n Diagram with Gust Lines for 60 …………………… 57
Figure 6.7: Mission 1 Profile……………………………………………… 59
Figure 6.8: Mission 2 Profile……………………………………………… 61
Figure 6.9: Mission 3 Profile………………………………………………. 63
xiii
LIST OF TABLES
Table 2.1: Balance Maximum Loads and Expected Balance Maximum Loads 12
Table 3.1: Results Summarization of Wing Billow Investigation………… 20
Table 3.2: Results Summarization of Reducing Wing Deformation …….. 23
Table 5.1: Results Summarization of Variation of Leading Edge Sweep… 32
Table 6.1: Flight Speed and Thrust Horsepower Required………………. 48
Table 6.2: Endurance at Cruise Velocities………………………………… 49
Table 6.3: Endurance at Maximum Velocities……………………………. 50
Table 6.4: Range at Cruise Velocities…………………………………….. 50
Table 6.5: Range at Maximum Velocities………………………………… 50
Table 6.6: Maximum Rate of Climb………………………………………. 52
Table 6.7: Minimum Glide Angle and Range……………………………. 53
Table 6.8: V-n diagram Velocities………………………………………… 54
Table 6.9: Turning Rates and Turning Radii for Cruise…………………… 58
Table 6.10: Turning Rates and Turning Radii for Maximum Velocity……... 58
Table 6.11: Mission 1 Summary…………………………………………….. 60
Table 6.12: Mission 2 Summary…………………………………………….. 62
Table 6.13: Mission 3 Summary…………………………………………….. 64
Table A.1: Run Log Series 1………………………………………………… 69
Table A.2: Run Log Series 2………………………………………………… 69
xiv Table B1: Experimental Results Series 1 – Effects of Billow………………. 72
Table B2: Experimental Results Series 1 – Effects of Reducing Wing Deformation 73
Table B.3: Experimental Results Series 2 – Effects of Leading Edge Sweep.. 74
Table B.4: Experimental Results Series 2 – Effects of Canard Deflection – Elevator 76
Table B.5: Experimental Results Series 2 – Effects of Canard Deflection – Aileron 77
Table B.6: Experimental Results Series 2 – Effects of Canard Deflection – Sideslip 78
xv
NOMENCLATURE
Parameters
AR Aspect Ratio C Coefficient K Alleviation Factor THP Thrust Horsepower R Turn Radius R Rate-of-Climb C S Wing Area SHP Shaft Horsepower U Velocity V Velocity W Weight b Span c Chord Length c Mean Aerodynamic Chord e Oswald Efficiency Factor g Gravity h Height k Thousand n Load Factor
Leading Edge Sweep Angle Billow Parameter Sideslip Angle Glide Angle Propeller Efficiency Mass Ratio Air Density Turning Rate
Subscripts
D Drag L Lift xvi M Pitching Moment U Velocity Y Side Force a Available de Derived Gust e Elevator g Gust l Rolling Moment max Maximum Value min Minimum Value n Yawing Moment o Zero Lift r Required
Slope Sideslip Angle a Aileron Deflection e Elevator Deflection Free Stream Conditions
xvii
CHAPTER 1
INTRODUCTION
Francis Rogallo pioneered the preliminary investigations into the aerodynamics of paragliders. The research was done to evaluate the practicality of a parawing design used for a reentry glider. The design included a highly swept wing made of two partial conical shaped surfaces and a light weight structure which supported a suspended load below the wing. This was controlled by tightening and loosening the cables connecting the wing and suspended load.
Rogallo’s preliminary low speed experiments were done with a 0.016-inch-sheet aluminum wing, while the high speed tests were done with the wing made of nylon parachute cloth. The results from the low speed tests showed that a wing made of non porous material performed well and could be flown at high angles of attack ranging from
20 to 90 . The results from the high speed tests showed none of the unfavorable characteristics of conventional parachutes, such as squidding and breathing. 1
William Sleeman Jr. and Joseph Johnson Jr. continued studies on the aerodynamics of parawings. Their investigations focused on the change in aerodynamics as the shape of the parawing was changed. They considered parawings which took the traditional conical shape and ones which took on a cylindrical shape. Again the
1 investigation was being applied to a reentry glider. Their investigations showed that a slight conical shape of the parawing resulted in the highest aerodynamic efficiency. 2
The advantages of a parawing are found in a light weight structure with a large area. This low wing loading has aerodynamic and performance advantages. Keeping the weight of the wings down allows for some flexibility of the vehicle design and propulsion system. Though much of the previous work on parawings has been directed towards reentry gliders, the major use of parawings has been in hang gliders, a sport that has developed from the original Rogallo wing. These air vehicles make use of another parawing attribute - packing. The wings may be folded for easy storage and transport.
One such application is the integration of a flexible parawing wing to an unmanned aerial vehicle, UAV, configuration. The concept incorporates a flexible parawing attached to an aerodynamic fuselage. This differs from Rogallo’s design of suspending a load below the wing, since the wing attaches directly to the fuselage. The concept raises several questions about the integration of the two: How is the performance of the wing affected from the air loads causing deformation in the wing shape? What happens to the longitudinal and lateral-directional stability and controllability now that the fuselage is no longer suspended below the wing? How does decreasing the leading edge sweep affect the performance of the parawing design, since the previous work concentrated on highly swept leading edges greater than 50 ?
The purpose of this thesis is to address these questions by two series of wind tunnel tests: the first to study the aerodynamics of a Rogallo type wing with a leading edge sweep of 30 , the second to evaluate the effects of sweep. The first test series uses a model of a wing structure which was capable of attaching a flexible material to it. The
2 wing surface material can be changed on a particular planform to create different deformations (or billow) due to the flight air loads. A second test series uses a model consisting of a flexible parawing mounted to an aerodynamic fuselage has been built and has the capabilities of leading edge sweep adjustments. The fuselage has added a control canard to provide longitudinal control. Leading edge sweeps are varied from 15 to 60 examine performance and potential morphing capabilities. Wind tunnel tests on these configurations to measure aerodynamic performance, longitudinal, and lateral-directional stability and controllability. Finally, as an example of flexible wing aerodynamics, the results are employed in the design of a small UAV that is contained in a 6” diameter by
36” tube.
3
CHAPTER 2
EXPERIMENTAL DESIGN / DEVELOPMENT / SETUP
2.1 Design Considerations
The Ohio State University Aeronautical and Astronautical Research Laboratory’s experience in aircraft innovations through experimental investigations was the foundation for developing the models and test matrices for the flexible wing aerodynamic study.
The first model used during the investigation consisted of a streamline fuselage with an externally mounted delta wing, which was designed to allow changes to the internal structure. The second model consisted of a larger fuselage, with canards for vehicle control, and an externally mounted delta wing. The wing was designed to allow changes to the internal structure of the wing as well as changes to the leading edge sweep angle.
Initial investigations determined the changes in the aerodynamic forces over a spectrum of different wing billows and changes to the internal structure of the wing. The aerodynamic effects due to changes in the leading edge sweep as well as canard deflections were also determined.
4 2.2 Experimental Facility
To conduct the Wind Tunnel Investigations, the AARL 3’ x 5’ Eiffel Type
Subsonic Wind Tunnel was selected as the test facility. The wind tunnel is capable of a maximum test section flow speed of 180 fps, produced by an 8 ft diameter 6-bladed fan that is belt driven with a 125hp 3 phase a.c. motor and variable frequency controller3
(Figure 2.1).
Figure 2.1: AARL 3’ x 5’ Eiffel Type Subsonic Wind Tunnel with Model Installed
The model was mounted inverted to a six component sting mounted balance.
Movement of the sting allowed for a range of angles of attack to be looked at during the investigation and keep the model centered in the tunnel.
The models were sized to reduce tunnel boundary or wall effects. In an open- circuit (Eiffel Type) wind tunnel, the angles of attack near the walls are drastically reduced and the effects become negligible by keeping the model span less than 80% of the tunnel width4. The model’s wing spanned covered a range of 46 inches to 28 inches,
77% to 47% of the tunnel width.
5 2.3 Wind Tunnel Model
2.3.1 Series 1 Wind Tunnel Model
For the first series of wind tunnel testing a model was built to represent the wing alone. The model was broken down into three components: wing structure, wing surface, and balance mount. Figure 2.2 shows a CAD drawing of the wind tunnel model.
Figure 2.2: Series 1 Wind Tunnel Model - 30
Aluminum was used to construct the wing structure. 0.125” thick sheet aluminum was used for the plate to mount the wing to balance. It measured 24” in length and the leading edges of the plate were cut at a 30 angle. The leading edge spars were made from 0.375” aluminum rods and measured 24” inches in length. A 7” slot was milled at one end of each spar for the mounting plate to slide into. Set screws were used to secure the leading edge spars and mounting plate together.
The wing surface was made from Kapton E polyimide film and has a thickness of
1.5 mils. The material properties of Kapton E include a tensile modulus 780,000 psi and a density of 0.0556 lbs . in3
6 The balance mount was made of 0.25” steel and welded into a triangular prism shape. The balance mounted inside of the hollow prism to allow for a flat surface to mount the wing.
2.3.2 Series 2 Wind Tunnel Model
For the second series of wind tunnel testing, a model was designed from the results obtained in the first series of wind tunnel testing. The model was broken down into four components: wing structure, wing surface, fuselage, and control canards. Figure
2.3 shows a CAD drawing of the wind tunnel model.
Figure 2.3: Series 2 Wind Tunnel Model - 30
A combination of stainless steel, aluminum, and plastic was used to construct the wing structure. Stainless steel rods were used for the leading edge wing spars and internal wing ribs. The leading edge spars measured 0.1875” in diameter and 24” in length, while the internal ribs measured 0.156” in diameter and 24” in length. A 1” strip of 0.125” thick sheet aluminum was used as a cross member that ran from the fuselage to the internal wing rib. A 0.125” thick sheet aluminum was used for the plate to mount the wing to fuselage and measured 24” in length. The spars and ribs were attached to the mounting plate by two 0.5” thick plastic mounts. Holes were drilled into the plastic 7 mounts to hold the spars, and ribs and were held in place with set screws. The mounts attached directly to the mounting plate and could be rotated to change the leading edge sweep. The wing surface was again made from Kapton E polyimide film and has a thickness of 1.5 mils.
The fuselage was constructed of fiberglass with a foam core. One layer of fiberglass was used for the model shell and Bond-O was used to fill in any imperfections during the curing process. It measured 31” in length from the nose to the tail. The balance was mounted inside of the fuselage.
The control canards were made of wood that had been shaped in an airfoil shape.
Each canard had a span length of 6” and a chord length of 2”. The canards mounted directly to the model fuselage.
8 2.4 Instrumentation
2.4.1 Hardware
A HH-388 multi-piece 6-component internal strain gauge force balance was used to measure forces during wind tunnel testing. The balance has an overall length of 8.99” with a maximum diameter of 1.125”. Forces and moments were measured using a combination of four strain gauges arranged into a Wheatstone bridge configuration. 5
The balance is shown in Figure 2.4.
Figure 2.4: Six Component Balance
The balance was powered from a Harrison Laboratories, Inc. Model 802B DC power supply. The output voltages were amplified with a Bay Laboratories, Inc. Model
5301 Differential Wideband DC amplifier and linked to a National Instruments Model
SCB-68 data acquisition device. 5 The power supply and amplifiers are shown in figure
2.5.
9
Figure 2.5: Power supply and Amplifiers 5
Previous work done by Rusincovitch , showed that the balance could be properly calibrated and produce accurate measurements during testing. Rusincovitch accredited the high level of accuracy in measurement despite the low loads, on the design of the balance and the high gain, low noise signal conditioning. For example in the case of the
10 normal force, the calibration process was completed by loading the balance at five locations acting in the normal direction. The weight ranged from 0 lbs to 20 lbs by 2 lbs increments and the voltage output was recorded. The process was completed for negative loading by rotating the balance 180 degrees and repeating the loading procedure. The slopes were then plotted for each loading position. Figure 2.6 shows the slopes per loading position. 5
Figure 2.6: Normal Force Calibration 5
A percent error between the measured loads and expected loads was found to be 0.027%.
This correlates to 0.0027lbs for an expected normal force of 10 lbs, and falls within the range such that the error does not exceed the measured loads. Table 2.1 shows the maximum design loads for the balance.
11 Table 2.1: Balance Maximum Loads and Expected Balance Maximum Loads 5
Expected Test % Component Max Load Load Maximum Normal Force 600 lbs 10 lbs 1.67% Side Force 600 lbs 3 lbs 0.5% Axial Force 150 lbs 1 lbs 0.67% Pitching Moment 1600 in-lbs 6 in-lbs 0.375% Yawing Moment 1600 in-lbs 3 in-lbs 0.187% Rolling Moment 400 in-lbs 1 in-lbs 0.25%
Balance measurements were taken in the body axis frame of reference. In this frame of reference the coordinate axis is fixed to the longitudinal axis of the aircraft. 6 To evaluate the aircraft performance and stability, the reference frame must be translated into the stability reference frame. Figure 2.6 shows the stability axis force components.
Figure 2.7: Stability Axis Force Components
The balance was mounted inside the tunnel using a sting. The system was designed to keep tunnel blockage at a minimum. With the model mounted inside of the tunnel an angle attack sweep of 0 to 32 was tested. The system contained a mounting bracket in which allowed the model to be placed at 5 and 10 side-slip. With the model
12 set at a side-slip configuration the angle of attack sweep was limited to a range of 0 to
20.
2.4.2 Data Acquisition and Reduction
A previously existing LabVIEW virtual instrument (VI) was used to record voltages from the balance. The VI took raw measurements taken by the balance in the body axis and translated the data to forces in the stability axis. User inputs describing the tunnel conditions, model configuration, and angle of attack allowed the transformation to take place. The data was saved to Excel spreadsheets after each wind tunnel run. The front panel of the virtual instrument is show in figure 2.7
Figure 2.8: LabVIEW Virtual Instrument Front Panel
MATLAB 7.1 was used to reduce the data. A script file read in data from each wind tunnel run and was used to produce the desired plots and charts.
13 2.5 Test Procedure
The first series of wind tunnel tests were started by taring the wing model through an angle of attack sweep. Once the balance was zeroed, the wind tunnel was run at testing conditions, and the wing model was run through an angle of attack sweep from
0 to 30 with measurements being recorded at two degree increments. The alpha sweep was then repeated for each wing billow configuration.
A rib was added to the wing model and an angle of attack sweep was completed to tare the model. The wind tunnel was run at the original testing conditions, and an angle of attack sweep was completed. A tare of the model was then completed, and an angle of attack sweep was completed for each addition of wing rib.
The second series of wind tunnel tests were started by taring the wing and fuselage model by running through an angle of attack sweep. The wind tunnel was run at the original testing condition, and an angle of attack sweep of 0 to 30 was completed for the model with no canard deflection. Data measurements were recorded at two degree increments. The process was then repeated two more times at different tunnel velocities.
The wind tunnel was then run at the original test conditions, and an angle attack sweep was completed for the model with an elevator deflection of the canard of -10 . An angle of attack sweep was then completed for elevator deflections of 5 , 10 , 15 , and
20 . Data measurements were again recorded at two degree increments.
The wind tunnel was then run at the original test conditions, and an angle attack sweep was completed for the model with an aileron deflection of the canard of -10 . An angle of attack sweep was then completed for aileron deflections of 5 , 10 , 15 , and
14 20 . Data measurements were again recorded at two degree increments. The model was then set at a -10 sideslip angle and an angle of attack sweep was completed for aileron deflections of -10 , 0 , 5 , and 10 .
The second series of wind tunnel testing was repeated for the four different wing leading edge sweeps.
15
CHAPTER 3
SERIES 1
FLEXIBLE WING INVESTIGATION AND RESULTS
The initial investigation focused on the effects of aerodynamic forces with changes in wing shape due to in flight air loads. A wing model having a leading edge sweep of 30 was tested. The change in the wing shape was modeled by adjusting the amount of wing billow created from the flexible wing material. A sheet of 1.5 mil
Kapton polyimide film was used to create the flexible wing surface. Further investigation focused on the effects of minimizing the flexible wing deformation to optimize performance. Figure 3.1 shows the wing model with half of the wing fixed with the
Kapton sheet and the other without to show the wing structure.
16
Figure 3.1: Wing model
The amount of billow tested during each run was defined using a ratio of one half the wing wetted area to the wing planform area. The ratio is defined as
0.5 S wetted (3.1) S planform
Wind tunnel runs for the first investigation were completed with =1.1, 1.05, 1.025, and
1.0 at a Reynolds number of 330,000. With =1.0, wing ribs were added to the wing support structure to reduce deformation. The investigation looked at wing configurations having 1, 3, and 7 ribs while runs were completed at the same Reynolds number.
17 3.1 Wind Tunnel Investigation
3.1.1 Effects of Billow
The effects of wing deformation were first established. The initial wind tunnel results showed a decrease in aerodynamic performance with greater amounts of deformation. It was observed that the flexible wing would vibrate significantly at low angles of attack and continued to vibrate until the wing was fully developed at higher angles of attack. The trailing edges of the wing would flutter violently throughout each wind tunnel run at all angles of attack. Figure 3.2 shows the lift performance for the initial investigation.
Lift Coefficient vs Angle of Attack =30 1.6
1.4
1.2
1
0.8
0.6 CL
0.4
0.2 =1.0 0 =1.025 =1.05 -0.2 =1.1
-0.4 0 5 10 15 20 25 30 AOA
Figure 3.2: Lift coefficient vs. Angle of attack with Wing Deformation
It was found that the aerodynamic performance would increase as the amount of deformation was decreased. With =1.1 the Kapton was observed to never fully develop, but vibrate violently throughout the complete angle of attack sweep. Wind tunnel runs with =1.0 produced the greatest maximum lift coefficient equal to 0.6.
With this configuration the wing had a standing wave that was observed to start at the
18 leading edge and end at the trailing edge of the wing. The trailing edge was again observed to flutter violently throughout the run.
The wind tunnel investigation produced similar L/D ratios as the billow was varied from one run to the next. The trend is visible when looking at the ratio of lift to drag also known as the aerodynamic efficiency. Figure 3.3 shows the ratio of lift to drag for the wing model with varied billow.
L/D vs Lift Coefficient =30 15
10
5
0 L/D
-5
=1.0 =1.025 -10 =1.05 =1.1
-15 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 CL
Figure 3.3: L/D vs. Lift Coefficient with Wing Deformation
Similar values of L/D were achieved throughout the investigation with maximum values near 6. The run with the greatest amount of billow showed large amounts of drag at similar lift coefficients due to the significant vibration of the Kapton wing which in turn gave low values of aerodynamic efficiency and is shown in the drag polar in figure 3.4.
19 Drag Polar =30 1.6
1.4
1.2
1
0.8 =1.0 =1.025 0.6 CL =1.05 0.4 =1.1
0.2
0
-0.2
-0.4 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 CD
Figure 3.4 Drag Polar with Wing Deformation
Table 3.1 summarizes the results found from the wind tunnel runs looking into the effects of wing billow.
Table 3.1: Results Summarization of Wing Billow Investigation
L CL C C D L max L L max max D 1.0 0.0375 0.61 6.93 0.25 1.025 0.0277 0.56 6.98 0.24 1.05 0.0287 0.59 5.7 0.28 1.1 0.0637 0.40 3.1 0.31
3.1.2 Effects of Reducing Wing Deformation
The effects of reducing wing deformation were developed from an addition of wing ribs between the leading edge of the wing and fuselage. Results produced from the addition of a single rib to the wing show a significant increase in aerodynamic performance. It was observed that the rib addition eliminated the violent flutter of the trailing edge. It was also observed that the traveling wave on the wing surface was eliminated. Figure 3.5 shows the wing model with the addition of 1 rib.
20
Figure 3.5: Wind Tunnel Model with 1 Rib
Further addition of ribs to the wing structure continued to improve the performance.
Runs completed with 3 and 7 ribs produced the highest values of lift coefficients, though similar values. It was concluded that reducing the deformation of the Kapton wing would improve performance to a point. After that point further addition of ribs was redundant.
Figure 3.6 shows the improved performance of lift with rib addition.
Lift Coefficient vs Angle of Attack =30 1.6
1.4
1.2
1
0.8
0.6 CL
0.4
0.2 0 ribs 0 1 ribs 3 ribs -0.2 7 ribs
-0.4 0 5 10 15 20 25 30 AOA
Figure 3.6: Lift Coefficient vs. Angle of Attack with Rib Addition
21 A CL max of 1.2 was achieved with model configuration having 3 and 7 ribs. The addition of ribs to the wing doubled the lift properties compared to the initial investigation.
Improvements were also gained with respect to the drag of the wing. At similar lift coefficients, the drag measurement was found to decrease with the addition of wing ribs and shown in figure 3.7.
Drag Polar =30 1.6
1.4
1.2
1
0.8 0 ribs 1 ribs 0.6 CL 3 ribs 0.4 7 ribs
0.2
0
-0.2
-0.4 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 CD
Figure: 3.7: Drag Polar with Rib Addition
Figure 3.8 shows the ratio of lift to drag for the wing with the addition of wing ribs.
22 L/D vs Lift Coefficient =30 15
10
5
0 L/D
-5
0 ribs 1 ribs -10 3 ribs 7 ribs
-15 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 CL
Figure 3.8: L/D vs. Lift Coefficient with Rib Addition
The lower amounts of drag and higher values of lift produced significantly better values of the aerodynamic efficiency. A lift to drag ratio near 15 was achieved with the models having 3 and 7 and ribs. These values more than doubled the aerodynamic efficiency found without any rib addition. Table 3.2 summarizes the results found from the wind tunnel runs looking into reducing wing deformation.
Table 3.2: Results Summarization of Reducing Wing Deformation
L CL C C D L max Ribs L L max max D 1.0 0 0.0375 0.61 6.39 0.25 1.0 1 0.0649 1.05 9.8 0.51 1.0 3 0.0772 1.17 14.13 0.66 1.0 7 0.0818 1.17 12.8 0.55
23
CHAPTER 4
THE UAV TODAY
The development of the UAV has become a major area of research in the aerospace community. Vehicles and technologies are being designed and researched to be applied to endless applications. The aerospace company Aerovironment has dedicated a large portion of the company’s resources into the development of UAVs. The product line incorporates small vehicles weighing only a few pounds to larger vehicles weighing a few thousand pounds.
Aerovironment has had success with the development of its small UAVs such as the Dragon Eye and the Raven. 8 The Dragon Eye is being used primarily by the US
Marines for reconnaissance and surveillance. The back-packable UAV is bungee- launched and weighs only six pounds with its 3’ 9” wingspan. The Raven on the other hand is being used by the US Army for low altitude reconnaissance and surveillance intelligence. Having a 4’ 6” wingspan and weighing four pounds the Raven is hand launched rather then launched with a rubber band.
Though these small UAVs give soldiers an advantage in the fight, the packaging size causes a hindrance to the soldier’s mobility. Integrating a flexible parawing to a
UAV configuration could significantly reduce the packaging size, without loosing any
24 aerodynamic performance. Other advantages in the parawing UAV is found in the possibility of morphing capabilities of the wing to adjust performance abilities best suited for mission requirements. That gives a broader spectrum of operation capabilities for a single UAV design all while reducing the vehicle weight from a lighter wing structure.
A problem arises with the integration of a flexible parawing to a UAV in the area of vehicle control. Having the inability to put control surfaces on the wing due to the flexible material limits the options for controlling the vehicle. A few solutions are possible to solve the control problem and include thrust vectoring and the addition of a control canard. Thrust vectoring adds unneeded complexity to the UAV design, while a control canard is easy to implement without adding complexity.
The second series of wind tunnel tests focus on the aerodynamic trends associated with changing the leading edge sweep of the flexible parawing and the stability and controllability with control canard deflections.
25
CHAPTER 5
SERIES 2
FLEXIBLE WING AND FUSELAGE INVESTIGATION AND RESULTS
Results from the “wing only” tests were applied to the investigation between the integration of the flexible Kapton parawing and vehicle fuselage. The focus of the investigation was on the effects of aerodynamic forces on a flexible parawing mounted directly to a vehicle fuselage while varying leading edge sweep. Also the longitudinal and lateral-directional stability and controllability from a control canard were investigated with varying leading edge sweep and sideslip angle.
The wind tunnel model was broken down into two parts; the vehicle fuselage with control canard and flexible parawing. The parawing design used during the wind tunnel investigation consisted of a sheet of 1.5 mil Kapton polyimide film with 1.0 for the flexible wing surface. While the parawing structure consisted of having one internal rib to reduce the deformation of the flexible surface and was mounted directly to the vehicle fuselage. The fuselage was made of fiberglass with mounts imbedded into it for the wing and two six inch canards.
The wind tunnel investigations were developed by installing the model into the
AARL 3’ x 5’ Eiffel Type Subsonic Wind Tunnel with varying tunnel conditions and
26 model configurations. Figure 5.1 shows the model mounted inside of the wind tunnel with a configuration having 30 .
Figure 5.1: Wing and Fuselage model
The wind tunnel investigation looked at leading edge sweeps of 15 to 60 at 15 increments with sideslip angles of 0 and -10 . The control canard acted as both an aileron and elevator, with a deflection range of 0 to 20 . The tunnel conditions were run at Reynolds numbers of 250, 330, and 400k at leading edge sweeps of 15 ,30 , and
45 . The configuration with a leading edge sweep of 60 was run at Reynolds numbers of 330, 400, and 500k.
27 5.1 Wind Tunnel Investigation
5.1.1 Effects of Leading Edge Sweep
The effects due to leading edge sweep were first established. A full wind tunnel test series was completed for each model configuration. Figure 5.2 shows the maximum
lift coefficient decreases with increasing leading edge sweep. A CL max of 1.51 was achieved with model configuration of 15 leading edge sweep.
Lift Coefficient vs Angle of Attack 1.6
1.4
1.2
1
0.8
0.6 CL
0.4
0.2 = 15 0 = 30 = 45 -0.2 = 60
-0.4 0 5 10 15 20 25 30 AOA
Figure 5.2: Lift coefficient vs. Angle of attack with Varied Leading Edge Sweep
Stall conditions were also found to appear at greater angles of attack for increasing sweep. This trend was expected for low aspect ratio delta wings due to the increased vortex lift acting on the wing. 7 This phenomenon is highly visible with model configurations of 60 sweep. The amount of vortex lift felt acting on the wing increases as the sweep is increased. This allowed the wing to stall at a very large angle of attack and achieve a moderate .
As the leading edge sweep of the wing is increased, surface area which contributes to the profile drag of the aircraft decreases. It was expected that an increase
28 in drag would be found with smaller sweep angles. Figure 5.3 shows the drag to increase for the same lift coefficient when leading edge sweep is increased.
Drag Polar 1.6
1.4
1.2
1
0.8 = 15 = 30 0.6 CL = 45 0.4 = 60
0.2
0
-0.2
-0.4 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 CD
Figure 5.3: Drag Polar with Varied Leading Edge Sweep
From the drag polar, the profile drag, C Do , of each vehicle was found by squaring the lift coefficient and plotting it against the drag coefficient. Then by fitting a first degree polynomial to the linear portion of the data the profile drag was found for each model
2 configuration. Figure 5.4 shows C L vs. CD in the linear portion with the first degree polynomial.
29 CL2 vs . CD
1.2
1
0.8
2 0.6 CL
0.4
0.2 = 15 = 30 0 = 45 = 60
-0.2 0 0.05 0.1 0.15 0.2 0.25 0.3 CD
2 Figure 5.4: C L vs. CD with Varied Leading Edge Sweep
Intuition tells us that the profile drag should decrease with increasing leading edge sweep.
Profile drag is due to skin friction and flow separation of the vehicle, and as the wing area is decreased and drag due to flow separation remains constant, the profile drag should
decrease. The opposite trend was found to occur with C Do , and the profile drag increased with increasing leading edge sweep. An analytical component buildup of agreed with the trend of increasing profile drag with increasing . Then from the definition of total drag, 9
C 2 C C L (5.1) D Do AR e the Oswald Efficiency Factor, e , was determined. It was found to decrease with increasing leading edge sweep. The values found for and e have been presented in table 5.1 on page 30.
The capabilities in the aerodynamic performance of the aircraft is dependent on the lift and drag in which it will encounter. It was shown that at a given angle of attack, 30 the lift increased with smaller leading edge sweep angles. When comparing lift and drag of the different model configurations, it was found that larger lift to drag ratios were found for smaller sweep angles, and the values decreased with increasing sweep. Figure
5.5 shows the ratio of lift to drag vs. lift coefficient.
L/D vs Lift Coefficient 15
10
5
0 L/D
-5
= 15 = 30 -10 = 45 = 60
-15 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 CL
Figure 5.5: L/D vs. Lift Coefficient with Varied Leading Edge Sweep
The flight conditions at which the maximum L/D ratios are found determine the cruise conditions for each model configuration. This point was found to occur near a lift coefficient of 0.4 and at an angle of attack near 8 .
The final aspect of aerodynamic performance on the longitudinal axis was to look at the pitching moment of the aircraft. This was done by looking at the pitching moment at different angles of attack with a static margin of -0.01 for all . A longitudinally stable aircraft will have a negative slope with respect to angle of attack. This stabilizes the aircraft due to disturbances, from equilibrium, or the trim condition. Figure 5.6 shows the pitching moment verse angle of attack.
31 Pitching Moment Coefficient vs Angle of Attack 0.1
= 15 0 = 30 = 45 = 60 -0.1
-0.2 Cm -0.3
-0.4
-0.5
0 5 10 15 20 25 30 AOA
Figure 5.6: Pitching Moment coefficient vs. angle of attack
Each configuration was found to have a stable trend with increasing angle of attack. The slope pertaining to each configuration was recorded for each UAV configuration. Table
5.1 summarizes the results found from the second series wind tunnel runs.
Table 5.1: Results Summarization of Variation of Leading Edge Sweep
L CL C C D L max C C L L max max D m Do e 15 0.1501 1.52 11.20 0.24 -0.0171 0.0060 0.79 30 0.0961 1.31 9.43 0.41 -0.0145 0.0080 0.61 45 0.0774 1.11 8.70 0.32 -0.0134 0.0140 0.44 60 0.0717 1.39 7.19 0.41 -0.0119 0.0160 0.37
5.1.2 Effects of Canard Deflection
The vehicle control characteristics were determined by measuring the effects of canard deflections. The first investigation looked into the effects of the canard acting as an elevator. The second determined the effects of the canard acting as an aileron.
Elevator deflections in general are used to provide longitudinal control of the aircraft. The convention used for positive elevator deflection with the model mounted
32 inverted inside the tunnel is when the elevator points up towards the tunnel ceiling. Then on the opposite hand a negative deflection is when the elevator points down towards the floor.
For a traditional aircraft the effect of a positive elevator deflection causes the aircraft to trim at a lower angle of attack. In the case of a canard, since the moment arms are switched from a traditional aircraft, a positive deflection causes the aircraft to trim at a larger angle of attack. It was found that this trend holds true for vehicle configurations with smaller leading edge sweep. Figures 5.7 and 5.8 show the change in pitching moment coefficient with changing elevator angle for 15 and 30 .
Pitching Moment Coefficient vs Angle of Attack =15 0.1 e = -10 e = 0 0 e = 5 e = 10 -0.1 e = 20
-0.2 Cm -0.3
-0.4
-0.5
0 5 10 15 20 25 30 AOA
Figure 5.7: Pitching Moment Coefficient vs. Angle of Attack ( 15 , for elevator deflection)
33 Pitching Moment Coefficient vs Angle of Attack =30 0.1 e = -10 e = 0 0 e = 5 e = 10 -0.1 e = 20
-0.2 Cm -0.3
-0.4
-0.5
0 5 10 15 20 25 30 AOA
Figure 5.8 Pitching Moment coefficient vs. angle of attack ( 30 , for elevator deflection)
For the lower angles of attack, a shift to increasing trim angles of attack is visible. As the sweep was increased, the interference caused from down wash of the canard increased. It was observed that the moment became unstable at larger angles of attack nearing vehicle stall and larger deflection angles. Figures 5.8 and 5.9 show the pitching moment coefficient with changing elevator angle for 45 and 60 .
34 Pitching Moment Coefficient vs Angle of Attack =45 0.1
0
-0.1
-0.2 Cm -0.3
-0.4 e = -10 e = 0 e = 5 -0.5 e = 10 e = 20
0 5 10 15 20 25 30 AOA
Figure 5.9: Pitching Moment coefficient vs. angle of attack ( 45 , for elevator deflection)
Pitching Moment Coefficient vs Angle of Attack =60 0.1
0
-0.1
-0.2 Cm -0.3
-0.4 e = -10 e = 0 e = 5 -0.5 e = 10 e = 20
0 5 10 15 20 25 30 AOA
Figure 5.10: Pitching Moment Coefficient vs. Angle of Attack ( 60 , for elevator deflection)
Canard deflections acting as ailerons as if to roll and yaw the aircraft, give an insight to the lateral-directional stability and controllability of each model configuration.
35 When considering the lateral-directional stability the rolling, and yawing moments about the center of gravity as well as the side force were looked into.
First the rolling moment was considered. It was measured with zero and positive ten degrees deflection. The convention used for positive aileron deflection with the model mounted inverted inside the tunnel was the right hand rule. It was observed that the rolling moment had little change between zero and ten degrees deflection for models with 15 , 30 , and 60 , but increased with the model configuration having 45 . This is shown in Figure 5.11.
Rolling Moment vs Aileron Deflection (a) 0.02
0.01
0
-0.01
-0.02 = 15 = 30
Cl -0.03 = 45 -0.04 = 60
-0.05
-0.06
-0.07
-0.08 -2 0 2 4 6 8 10 12 a
Figure 5.11: Rolling Moment for Aileron Deflection
The side force was observed to increase with aileron deflection, and is shown in figure
5.12.
36 Side Force vs Aileron Deflection (a) 0.12
= 15 = 30 0.1 = 45 = 60 0.08
0.06 CY
0.04
0.02
0 -2 0 2 4 6 8 10 12 a
Figure 5.12 Side Force vs. Aileron Deflection
The yawing moment was corrected for the center of gravity shift between model leading edge sweep angles. This was done since the static margin for the pitching moment was held constant at -0.1. An increase in yawing moment was observed with increasing positive elevator deflection. This is shown in figure 5.13.
Yawing Moment vs Aileron Deflection (a)
0.25
0.2
0.15
0.1
0.05
0 Cn
-0.05
-0.1
-0.15 = 15 -0.2 = 30 = 45 -0.25 = 60
-10 -5 0 5 10 15 20 a
Figure 5.13: Yawing Moment for Aileron Deflection
37 5.1.3 Effects Due to Sideslip
The effects of -10 sideslip were established for model configurations having
30 and greater. The model configuration with 15 was not tested because the wing span was too large to fit inside of the 3’ x 5’ test section once positioned at -10 sideslip. The yawing moment was again corrected to accommodate for center of gravity shift. Figure 5.14 shows the yawing moment effects due to sideslip.
Yawing Moment vs Sideslip ()
0.1
0
-0.1
-0.2 Cn
-0.3
-0.4 = 30 -0.5 = 45 = 60
-12 -10 -8 -6 -4 -2 0 2
Figure 5.14: Yawing Moment vs. Sideslip
The yawing moment was observed to increase between sideslip angles. This denotes a positive slope and a stable aircraft.
The rolling moment was observed to decrease between yaw angles. A slight decrease in the moment was observed for model configurations with leading edge sweeps of 30 and 60 . This is shown in figure 5.15.
38 Rolling Moment vs Sideslip () 0.02
0
-0.02 = 30 = 45 = 60
Cl -0.04
-0.06
-0.08
-0.1 -12 -10 -8 -6 -4 -2 0 2
Figure 5.15: Rolling Moment vs. Sideslip
The side was force was observed to increase between sideslip angles. Figure 5.16 shows the side force effects from sideslip.
Side Force vs Sideslip ()
0.05
0
-0.05 CY -0.1 = 30 = 45 = 60 -0.15
-0.2
-0.25 -12 -10 -8 -6 -4 -2 0 2
Figure 5.16: Side Force vs. Sideslip
The slopes pertaining to each moment and force for each model configuration were recorded and presented in the UAV configuration section of this work.
39 5.2 UAV Configurations
The goals of the two series of wind tunnel investigations were to look into the practicality of integrating a UAV with a flexible Kapton parawing and design a UAV capable of being contained in a 6” diameter by 36” tube. The results allowed for the aerodynamic, longitudinal and lateral-directional stability and controllability coefficients to be determined for four UAV configurations with varying leading edge sweep. A summarization of the coefficients as well as the three-view drawings and geometry of each vehicle are shown in figure 5.17 through 5.20.
The mean aerodynamic chord for each UAV configuration was found from equation (5.2). 9
2 1 2 c cr (5.2) 3 1
The taper ratio, , is defined as: 9
c t (5.3) cr
With each wing configuration being a delta wing, the taper ratio was effectively zero.
This in turn caused each wing to have the same mean aerodynamic chord of two thirds the root chord length.
40 5.2.1 15 Leading Edge Sweep
C C C C C C C C SeaLevel CL CD L D m LU DU mU Le me 15 0.5748 0.0514 0.1501 0.0227 -0.0171 -0.0014 -0.0001 0.0007 -0.0016 0.0032
CY Cn Cl C C C SeaLevel Ya na la 15 - - - 0.0043 0.0106 0.0002
Reference Geometry
41 2 S 5.24 ft c 1.33 ft b 4.16 ft AR 3.30
Figure 5.17 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle Geometry - Leading Edge Sweep
41 5.2.2 30 Leading Edge Sweep
C C C C C C C C SeaLevel CL CD L D m LU DU mU Le me 30 0.4136 0.0439 0.0961 0.0081 -0.0145 -0.0015 -0.0001 0.0003 -0.0049 0.0033
SeaLevel CY Cn Cl C C C Ya na la 30 0.0242 0.0508 -0.0003 0.0059 0.0145 -0.0001
Reference Geometry
42 2 S 4.17 ft c 1.33 ft b 3.63 ft AR 3.16
Figure 5.18 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle Geometry - 30 Leading Edge Sweep
42 5.2.3 45 Leading Edge Sweep
C C C C C C C C SeaLevel CL CD L D m LU DU mU Le me 45 0.3287 0.0377 0.0774 0.0077 -0.0134 -0.0015 -0.0001 0.0006 0.0072 0.0077
SeaLevel CY Cn Cl C C C Ya na la 45 0.0130 0.0274 0.0079 -0.0080 0.0234 0.0036
Reference Geometry
43 2 S 3.14 ft c 1.33 ft b 3.32 ft AR 3.51
Figure 5.19 UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle Geometry - 45 Leading Edge Sweep
43 5.2.4 60 Leading Edge Sweep
C C C C C C C C SeaLevel CL CD L D m LU DU mU Le me 60 0.4150 0.0577 0.0717 0.0095 -0.0119 -0.0003 0.0000 0.0002 0.0008 0.0147
CY Cn Cl C C C SeaLevel Ya na la 60 0.0223 0.0296 -0.0009 0.0063 0.0094 0.0006
Reference Geometry
44 2 S 2.08 ft c 1.33 ft b 2.63 ft AR 3.33
Figure 5.20UAV Aerodynamic, Stability, Controllability Coefficients and Vehicle Geometry - 60 Leading Edge Sweep
44
CHAPTER 6
UAV PERFORMANCE ANALYSIS
From the two series of wind tunnel tests, four UAV configurations which use a flexible Kapton wing were developed. The coefficients and parameters which were found for each configuration during the wind tunnel testing have been used to complete a performance analysis of each vehicle. The analysis was completed for sea level conditions and the weight of the Dragon Eye UAV, 6 lbs, was used. This weight was chosen due to the comparable size between the Dragon Eye and the group of UAV configurations. The analysis includes investigating the Thrust Horsepower Required and
Available, Range and endurance, Rates of Climb, Gliding Flight, V-n diagrams, and
Turning Flight. The results found during the performance analysis were used to describe a possible application of the UAV configurations.
A possible application for such a UAV is to complete a chemical assessment of the air to ensure a safe environment. Three possible scenarios in which a chemical assessment is needed has been described in detail, this includes stating the situation and the important guidelines for a successful mission. For each scenario an optimum UAV configuration was chosen from the characteristics found during the performance analysis to best complete the mission. The configuration remained constant throughout each
45 mission, meaning in flight morphing of the wing was neglected. The details of completing each mission which includes the elapsed time, distance traveled, and powered flight time remaining for each phase has been described and tabulated for each scenario.
46 6.1 Performance
6.1.1 Thrust Horsepower Required
For an aircraft to maintain level flight a certain amount of power is required to keep the aircraft aloft. The power required is dependant on a number of factors including wing area, air density, lift and drag coefficient and wing loading. The thrust horsepower required for level flight was determined from equation (6.1). 9
2 V CDoS 2 2 W 1 THPr V 2 (6.1) 550 2 e b V with b as the wing span, e as the Oswald Efficiency Factor, and S being the wing area.
Values for C Do were determined from figure 5.4, while was determined using the definition of total drag. The values for and are tabulated in table 5.1. Figure 6.1 shows the thrust horsepower required for each UAV design.
THP Required vs. Velocity 0.5 = 15 0.45 = 30 = 45 0.4 = 60 0.35 Available THP
0.3
0.25 THP
0.2
0.15
0.1
0.05
0 0 10 20 30 40 50 60 70 80 90 100 V knots
Figure 6.1: Thrust Horsepower Required
The thrust horsepower gives an insight to the flight speed regime that the aircraft is capable of at a given altitude. The limiting factor to the flight speed is the available
47 power given from the engine. To achieve greater velocities an engine capable of producing more thrust is needed. Table 6.1 shows the important flight speeds and the thrust horsepower required to achieve it. The velocities are in knots.
Table 6.1: Flight Speed and Thrust Horsepower Required
Vstall THPrstall Vcruise THPrcruise Vmax THPr max 15 15 0.0563 25 0.0477 85 0.2125 30 18 0.0613 32 0.058 82 0.2125 45 23 0.061 41 0.0671 74 0.2125 60 26 0.0811 46 0.0715 77 0.2125
6.1.2 Thrust Horsepower Available
In order for an aircraft to maintain level flight the engine must be capable of producing the need power. For propeller driven aircraft, the available power is dependent on the shaft horsepower and propeller efficiency. 9
THPA SHP (6.2)
A Hacker A30-12m electric motor was chosen to power each UAV. The motor is capable of producing 350 watts at 40 amps and 12.5 V. The motor was rated for cruise at
22 amps and 7.7 V. Using a linear interpolation between the max power and the rated cruise, it was determined the motor produced 192.5 watts of power. This converted to
0.26 shaft hp.
To power the motor a lithium-ion polymer battery was chosen. A battery producing 11.1 V with 5000 mAh was originally chosen for the power supply. On further investigation of the battery life with the Hacker motor, the flight time was limited to 45 minutes. To increase the flight time, three batteries were connected in parallel to triple the mAh to 1500 while holding the voltage at 11.1V.
48 To determine the thrust horsepower available, a generic propeller efficiency across the speed regime was used. The available thrust horsepower is shown in figure
6.1.
6.1.3 Range and Endurance
With traditional gas powered engines the defining parameter of the range and endurance is the specific fuel consumption. With electric power vehicles the defining parameter of range and endurance is the battery amp hour. The amp hour determines how long a battery can sustain a continuous amp draw over a period of time. For example a battery rated for 5 Ah can withstand a draw of 5 amps for one hour.
The battery chosen to power the electric motor is rated at 5Ah. To increase the flight time 3 batteries were connected in parallel to increase the rating to 15Ah and not increase the voltage. The range and endurance were predicted for cruise and maximum velocities.
To calculate the endurance, the THPrcruise was converted to watts and then the amps were determined from an interpolation between the watts and amperage from the motor ratings. The flight time was then calculated by dividing the amp hours by the amperage pull. Tables 6.2 and 6.3 show the endurances for each UAV at cruise and maximum velocities.
Table 6.2: Endurance at Cruise Velocities
Endurance THPr Watts Amps (mins) 15 0.0477 35 4 225 30 0.058 43 5 180 45 0.0671 50 5.7 157 60 0.0715 53 6 150
49 Table 6.3: Endurance at Maximum Velocities
Endurance THPr Watts Amps (mins) 15 0.2125 158.46 18.1 49 30 0.2125 158.46 18.1 49 45 0.2125 158.46 18.1 49 60 0.2125 158.46 18.1 49
The range was determined for each UAV for both cruise and maximum velocities.
Although the UAVs with larger wing area are able to stay in the air longer, they have the shortest range. Tables 6.4 and 6.5 show the ranges for each UAV at cruise and maximum velocities.
Table 6.4: Range at Cruise Velocities
V V Endurance Nautical (knots) ( ft / sec) (mins) Miles Miles 15 24.3 41.01 225 104.86 91.11 30 32 54.01 180 110.47 95.99 45 41 69.2 157 123.45 107.27 60 45.3 76.46 150 130.32 113.23
Table 6.5: Range at Maximum Velocities
(mins) 15 85 143.46 49 79.88 69.41 30 82 138.40 49 77.06 66.96 45 74 124.90 49 69.55 60.43 60 77 129.96 49 72.36 62.88
The maximum endurance at cruise was found to be near four hours, while the minimum endurance was found to be two and a half hours. To improve the endurance on the aircraft, additional batteries would be needed and connected in series with the existing
50 cell. The maximum range at cruise was found to 130 miles at 45 knots, while a minimum range was found to 104 miles at 24 knots.
6.1.4 Rate of Climb
The power required and power available predictions previously discussed in section 6.1.1 and 6.1.2 give an insight to the climb performance of each UAV configuration. As the power output from the motor is increased beyond the amount that is needed for level flight, the aircraft will begin to climb. If the difference between what is needed and what is available is increased, the rate at which the aircraft will climb will also increase. The flight speeds in which the available power is greater then the required, denotes where climbing is possible. This is shown in figure 6.1.
The velocities at which the maximum rate of climb is possible, correspond to the velocities with the greatest amount of excess power. This point is denoted as the apex of the plot looking at rate of climb for a given velocity. The rate of climb was found from using equation (6.3). 9
33000(THP THP ) R a r (6.3) C W
With Rate-of-Climb in ft . Figure 6.2 shows the sea level rates of climb per velocity min for each UAV design.
51 R/C vs Air speed 1000
900
800
700
600
500
400 R/C ft/minR/C 300
200 R/C =15 100 R/C =30 R/C =45 0 R/C =60
-100 0 50 100 150 ft/sec
Figure 6.2: Sea Level Rate of Climb
The maximum rate of climb was found to decrease with increasing leading edge sweep and to occur at greater velocities for great values of . Table 6.6 shows the maximum rate of climbs for each UAV design.
Table 6.6: Maximum Rate of Climb
V ft R Cmax sec 15 878 63 30 792 66 45 624 69 60 450 82
With a cruising altitude of 1000 feet, the time to climb to altitude ranges from less then a minute for 15 to two and a half minutes for 60
6.1.5 Gliding Flight
For some applications of the UAV configuration, a scenario will occur where the battery life is exceeded with a resulting total loss of power. For these particular scenarios,
52 the gliding characteristics of the aircraft are important to see how the UAV will perform under these conditions. To achieve the largest range, the aircraft needs to fly at the lift coefficient corresponding to the maximum lift to drag ratio. The minimum glide angle is also required to maximize the range and can be determined from equation (6.4). 10
1 1 min tan (6.4) L Dmax
In the cases for the four UAVs as the sweep is increased, the maximum lift to drag ratio decreases. As the lift to drag ratio decreases, the glide angle increases and reduces the range of the steady glide. Table 6.7 shows the minimum glide angle to achieve the maximum range for each UAV design.
Table 6.7: Minimum Glide Angle and Range
Range L Dmax min (1000 ft) 15 11.80 4.8 2.23 30 9.42 6.1 1.78 45 8.70 6.6 1.65 60 7.19 7.9 1.36
The range was determined from equation (6.5). 10
L Rmax h (6.5) Dmax
With the units of range in miles.
6.1.6 V-n Diagrams
The structural design of the aircraft is an important factor in the design process.
The first step in a structural analysis is to look at the plot of the maximum loading factor
53 for a given flight velocity. This is know as the V-n diagram and shows the loading that will either stall the aircraft or cause catastrophic damage to the aircraft.
Typical maneuver load limits for a general aviation aerobatic aircraft are set at
6 and 3. 7 The maneuver limits for the UAVs have been set at 4 and 3 . The
10 bound defining the stall region is a function of CL max and defined as:
1 C n V 2 L max (6.6) max 2 W S
The dive velocity, the maximum flight velocity boundary is defined so the cruise velocity is not greater then 80% of it. In this case it was defined as 20% greater then the maximum velocity. The important velocities for each V-n diagram are listed in table 6.8.
While the vehicle with 15 produced a design velocity for maximum gust of
100 ft . sec
Table 6.8: V-n diagram Velocities
V Design VCruise VDive 15 51 41 171 30 61 54 165 45 77 69 148 60 87 76 162
During flight an aircraft will encounter gusts. Depending on the strength of the gust, it may cause the aircraft to exceed the maneuver load limits and essentially stall the aircraft. The standard gust, U , that an aircraft can handle is 30 ft . 10 A gust line for de sec a given gust velocity is defined in equation (6.7). 9
K V U C n 1 g E de L (6.7) 2W S
54 With CL as the lift curve slope, the flight speed, VE , and the Alleviation Factor, K g , defined as
0.88 g K g (6.8) 5.3 g
and the mass ratio, g , defined as
2W S g (6.9) c CL g
The gust lines were added to the V-n diagram and show if changes to the load limits are needed. 11 Figures 6.3 through 6.6 show the V-n diagrams with 30 ft gust lines for sec each UAV configuration.
V-n =15 5
4
3
2 Gust at 30 ft/sec V Cruise
1 n 0
-1
-2
-3
-4 0 20 40 60 80 100 120 140 160 180 V ft/sec
Figure 6.3: V-n Diagram with Gust Lines for 15
55 V-n =30 5
4
3
2 Gust at 30 ft/sec
1 V Cruise n 0
-1
-2
-3
-4 0 20 40 60 80 100 120 140 160 180 V ft/sec
Figure 6.4: V-n Diagram with Gust Lines for 30
V-n =45 5
4
3 Gust at 30 ft/sec V Cruise 2
1 n 0
-1
-2
-3
-4 0 20 40 60 80 100 120 140 160 180 V ft/sec
Figure 6.5: V-n Diagram with Gust Lines for 45
56 V-n =60 5
4
3
2
1 n 0
-1 Gust at 30 ft/sec -2 V Cruise
-3
-4 0 20 40 60 80 100 120 140 160 180 V ft/sec
Figure 6.6: V-n Diagram with Gust Lines for 60
6.1.7 Level Turns
Level turns for an aircraft are highly dependant on the load factor in which the aircraft can withstand and the amount of lift produced. Now that the loading limits for each UAV have been defined, the turning radius and rates of each can be investigated.
For aircraft the turning radius and rate are defined as 10
V 2 R (6.10) g n 2 1 and
g n 2 1 (6.11) V with n representing the load factor. This then tells us dynamics of the aircraft during a turning maneuver, then aids in determining mission requirements. With having the lift and loading factor limits defined, the minimum turning radius and maximum turning rate
57 can be found. Table 6.9 and 6.10 show the minimum turning radii and maximum turning rates for each UAV design for both cruise and maximum velocities.
Table 6.9: Turning Rates and Turning Radii for Cruise
C n V ( fps) L max W S max Rmin 15 2.6 41.01 1.52 6 5.24 2.06 19.68 30 3.1 54.01 1.3 6 4.17 1.86 28.91 45 3.4 69.20 1.1 6 3.14 1.55 45.37 60 3.1 76.46 1.3 6 2.08 1.31 57.96
Table 6.10: Turning Rates and Turning Radii for Maximum Velocity
15 4 143.46 1.52 6 5.24 2.56 19.68 30 4 138.40 1.3 6 4.17 2.11 28.91 45 4 124.90 1.1 6 3.14 1.68 45.37 60 4 129.96 1.3 6 2.08 1.49 57.96
9 With Rmin and max defined in equations (6.12) and (6.13).
2 W Rmin (6.12) gCL max S and
C n g L max max (6.13) max 2W S
58 6.2 Mission Scenarios and Profiles
6.2.1 Mission 1
A biological attack has taken place on Washington D.C. and the city has been evacuated for a 50 mile radius. The President is held up inside a bunker in Washington and is waiting for safe evacuation. Armed forces stationed in Baltimore 35 miles away need to assess the chemical composition of the air to secure a safe evacuation of the president. A 60 min analysis of the air is required for an accurate assessment.
To complete the mission, the UAV configuration with 15 was chosen. It was chosen due to the greatest endurance, capable of handling the long loiter time needed for air evaluation. The mission profile was broken down into four sections. These include climb, cruise, loiter, and return. Figure 6.7 gives a visual representation of the mission.
Figure 6.7: Mission 1 Profile
During the climb phase of the mission the vehicle is capable of a maximum rate of climb equal to 878 ft at 63 ft . The cruise altitude for the mission is 1000 ft min sec and travels 0.80 miles during the climb phase in a little over one minute. Once the cruise phase is started the aircraft needs to now travel 34.2 miles to Washington. Traveling at
41 it takes 74 minutes to reach Washington.
59 For the loitering phase of the mission, the aircraft makes figure eight maneuvers with a turn radius of 20 ft. This is the minimum turning radius for cruise velocity.
During the 60 minute loiter time, the aircraft will make a total 2,349 half turns and cover a distance of 28 miles.
During the return flight, the vehicle will travel at 41 ft for 35 miles, taking 75 sec minutes. The total flight time for the mission is 210 minutes and traveling 98 miles.
Which falls under the range and endurance of the UAV with 15 . A summary of the mission is presented in table 6.11.
Table 6.11: Mission 1 Summary
Time Distance Flight Time Remaining V Phase (mins) (mi) (mins) (fps) 1 0 0 225 0 2 1 0.8 224 63 3 74 34.20 150 41 4 60 28 90 41 5 75 35 15 41 End 210 98 15 0
6.2.2 Mission 2
Soldiers are in the process of launching a ground attack on a city in Iraq and are scheduled to strike in 12 minutes. Late intelligence from the city has been acquired from base command, 14.5 miles to the south of the city, that an anthrax contamination of the target area has taken place. Base command needs to give the final go or no go authorization before the strike is to begin. A 2 minute analysis of the air is required for an accurate assessment.
To complete the mission, the UAV design with 30 was chosen. It was chosen due to the high flight speed needed for a time sensitive mission even though the
60 UAV design with 15 has a higher maximum flight velocity. At that flight velocity a
30 ft gust will push the UAV past the positive load limit. The mission profile was sec broken down into four sections. These include climb, cruise, loiter, and return. Figure
6.8 gives a visual representation of the mission.
Figure 6.8: Mission 2 Profile
During the climb phase of the mission the vehicle is capable of a maximum rate of climb equal to 792 ft at 66 . The cruise altitude for the mission is 1000 ft min and travels 0.95 miles during the climb phase in 1 minute and 14 seconds. Once the cruise phase is started the aircraft is required travel 13.55 miles in less than 8 minutes and
46 seconds. Traveling at 138 , it takes 8 minutes and 38 seconds. The total flight time to reach the combat zone is 9 minutes and 42 seconds. Once the analysis is complete, a 18 second window is available to give authorization.
Once in the loitering phase of the mission, the flight speed is dropped to the cruise velocity for the UAV, 54 . A figure eight maneuver, with a turn radius of 28.9 ft is performed during the analysis. This is the minimum turning radius for cruise velocity.
61 During the 2 minute loiter time, the aircraft will make a total of 35.6 full turns and cover a distance of 1.23 miles.
For the return flight, the UAV travels at the cruise velocity of 54 ft and takes sec a total of 23 minutes and 38 seconds. The total flight time for the mission is 26 minutes and 46 seconds and traveling 25.57 miles. A summary of the mission is presented in table 6.12.
Table 6.12: Mission 2 Summary
Time Distance Flight Time Remaining V Phase (mins) (mi) (mins) (fps) 1 0 0 49 0 2 1.30 0.95 47.70 68 3 8.76 13.55 38.94 121 4 2 1.23 36.94 69 5 23.63 14.5 13.31 69 End 35.69 25.57 13.31 0
6.2.3 Mission 3
A nuclear power company has just endured simultaneous reactor failures at their nuclear plants located in Oak Harbor, Ohio, 94 miles away in Mentor, Ohio. The company’s engineers are in the process of determining the severity of the contamination.
They are stationed at a facility located 10 miles outside of Oak Harbor. A 15 mph wind is carrying the fallout from Oak Harbor towards Mentor. A 15 min analysis of the air is required for an accurate assessment.
To complete the mission, the UAV design with 60 was chosen. It was picked due to the large range needed. The mission profile was broken down into seven sections. Including multiple phases of cruise and loiter, with a climb and glide phase.
Figure 6.9 gives a visual representation of the mission.
62
Figure 6.9: Mission 3 Profile
During the climb phase of the mission the vehicle is capable of a maximum rate of climb equal to 450 ft at 82 ft . The cruise altitude for the mission is 1000 ft. min sec
The vehicle travels 1.89 miles during the climb phase in 2 minutes and 29 seconds. The initial cruise to the Oak Harbor nuclear plant takes 9 minutes and 23 seconds, traveling at
76 . Once the UAV reaches the plant, it begins a figure eight maneuver to complete the air analysis. The flight path takes a turn radius of 58 ft and makes 189 turns, traveling
13 miles during the 15 minute analysis time.
The second cruise phase of the mission from Oak Harbor to Mentor involves a
15mph tail wind. The tail wind adds 22 to the ground speed of the UAV, pushing it to 98 . This allows the 94 mile trip to be completed in 1 hour and 41 minutes.
The second loitering phase is completed in the same fashion as the first. A figure eight maneuver covering 13 miles is performed for the analysis. At this point in the mission, the flight time is at 127 minutes. The maximum endurance for the UAV is 150 minutes. This gives 23 minutes of flight time left before the batteries are drained.
After the second loitering phase of the mission, the UAV changes it’s heading to the south towards Solon, Ohio, 21.5 miles away. With 23 minutes of flight time, the
63 UAV can travel 19.9 miles at 76 ft before the batteries are depleted. With 1.6 miles sec left before Solon is reached, the UAV begins the gliding flight phase of the mission.
With the UAV positioned to achieve the maximum lift to drag ratio, the minimum glide angle can be obtained to maximize the range. At a minimum glide angle of 7.9 , the
UAV can glide for 1.36 miles from the 1000ft cruise altitude before touching down roughly a quarter mile outside Solon. A summary of the mission is presented in table
6.13.
Table 6.13: Mission 3 Summary
Time Distance Flight Time Remaining V Phase (mins) (mi) (mins) (fps) 1 0 0 150 0 2 2.48 1.89 147.52 74 3 9.39 8.11 138.13 76 4 15 13 123.13 76 5 84.4 94 38.73 94 6 15 13 23.73 76 7 23.95 20 0 76 8 1.57 1.36 0 0 End 151.57 151.36 0 0
64
CHAPETER 7
SUMMARY AND CONCLUSION
The wind tunnel investigations were done in order to understand the aerodynamic performance of a flexible parawing and to how it pertains to a UAV application. The analysis was completed at The Ohio State University Aeronautical and Astronautical
Research Laboratory facility with the use of the 3’ x 5’ Eiffel Type Subsonic Wind
Tunnel.
One investigation included wind tunnel testing of a flexible Kapton parawing with
30 leading edge sweep delta wing as the amount of wing billow is decreased. The wing billow was defined as a ratio of one half the wetted area to the planform area. Results
from the wind tunnel runs produced a CL max 0.6 with a billow parameter equal to one.
An additional wind tunnel investigation of the wing looked at the effects of increasing the structural soundness of the flexible parawing wing structure to optimize aerodynamic performance. This was done by the addition of wing ribs between the leading edge and fuselage. Results from the investigation that the addition of wing ribs
increased the aerodynamic performance and a CL max 1.2 was obtained with model configurations with 3 and 7 ribs per wing.
65 The final wind tunnel investigation focused on an integration between the optimized wing design, fuselage shape, and canard design. The testing of the UAV system included runs to analyze the aerodynamic performance as well as the longitudinal and lateral-directional stability as the leading edge sweep was varied from 15 to 60 by
15 increments. Results from the investigation showed CL max values to decrease with
decreasing leading edge sweep from 15 to 45 . A CL max 1.51 was obtained with
15 . At a leading edge sweep of 60 , the contributing vortex lift at higher angles of
attack allowed a CL max 1.3 to be obtained.
Longitudinal and lateral-directional stability coefficients were found at cruise conditions for each model configuration from the data set and used with the aerodynamic analysis to predict the respective UAV performance. The performance predictions included topics on thrust horsepower, range, endurance, aircraft structure, turning flight, climbing flight, and gliding flight.
From the performance predictions, three possible mission scenarios were described and a full mission profile was created for each mission. The vehicle configuration that was chosen to complete each mission was based on key performance abilities of the vehicle pertaining to each respective mission.
66
REFERENCES
1. Rogallo, Francsis M. Preliminary Investigation of a Paraglider. National Aeronautics and Space Administration. NASA TN D-443. August 1960
2. Sleeman Jr., William C. and Johnson Jr., Joseph L. Aerodynamics of the Parawing. AIAA Magazine Article, June 1963
3. Janiszewska, Jolanta M. Three Dimensional Aerodynamics of a Simple Wing in Oscillation Including Effects of Vortex Generators. The Ohio State University, 2004.
4. Pope, Alan. Low Speed Wind Tunnel Testing. 1st Edition. John Wiley & Sons Inc. 1966.
5. Rusincovitch, Jeff. On The Development of a Six Component Platform For Testing Aircraft Models. The Ohio State University. 2004.
6. Etkin, Bernard. Dynamics of Flight – Stability and Control, 2 nd Edition. John Wiley & Sons, Inc. 1982
7. Torres, Gabriel and Mueller, Thomas. Low-Aspect-Ratio Wing Aerodynamics at Low Reynolds Numbers. AIAA Journal, Vol.42, No. 5, pp.865-873
8. Aerovironment. 1 Jan. 2009. Aerovironment. 27 Feb. 2009 www.avinc.com.
9. Gregorek, Gerald M. Aircraft Performance. AAE 200/201 Notes. 2005
10. Anderson Jr., John D. Introduction to Flight. 5 th Edition. The McGraw-Hill Companies, Inc. 2005
11. Raymer, Daniel P. Aircraft Design: A Conceptual Approach. 4 th Edition. American Institute of Aeronautics and Astronautics, Inc. 2006
67
APPENDIX A
RUN LOGS
68 Table A.1: Run Log Series 1
Run Date Title Ribs 0 1 3 7 1.00 1.025 1.05 1.01 1 3/13/2008 Series 1 - Wing x x 2 3/13/2008 Series 1 - Wing x x 3 3/13/2008 Series 1 - Wing x x 4 3/13/2008 Series 1 - Wing x x 5 5/21/2008 Series 1 - Wing x x 6 5/22/2008 Series 1 - Wing x x 7 5/23/2008 Series 1 - Wing x x
Table A.2: Run Log Series 2
Run Date Title e a Re 15 30 45 60 1 9/18/2008 Series 2 - Wing - Fuselage x 0 0 0 250 2 9/18/2008 Series 2 - Wing - Fuselage x 0 0 0 330 3 9/18/2008 Series 2 - Wing - Fuselage x 0 0 0 400 4 9/18/2008 Series 2 - Wing - Fuselage x 5 0 0 330 5 9/18/2008 Series 2 - Wing - Fuselage x 10 0 0 330 6 9/18/2008 Series 2 - Wing - Fuselage x 20 0 0 330 7 9/18/2008 Series 2 - Wing - Fuselage x -10 0 0 330 8 9/18/2008 Series 2 - Wing - Fuselage x 0 -10 0 330 9 9/18/2008 Series 2 - Wing - Fuselage x 0 5 0 330 10 9/18/2008 Series 2 - Wing - Fuselage x 0 10 0 330 11 9/18/2008 Series 2 - Wing - Fuselage x 0 20 0 330 12 9/19/2008 Series 2 - Wing - Fuselage x 0 0 -10 330 13 9/19/2008 Series 2 - Wing - Fuselage x 0 -5 -10 330 14 9/19/2008 Series 2 - Wing - Fuselage x 0 -10 -10 330 15 9/19/2008 Series 2 - Wing - Fuselage x 0 5 -10 330 16 9/19/2008 Series 2 - Wing - Fuselage x 0 10 -10 330 17 9/22/2008 Series 2 - Wing - Fuselage x 0 0 0 250 18 9/22/2009 Series 2 - Wing - Fuselage x 0 0 0 330 19 9/22/2010 Series 2 - Wing - Fuselage x 0 0 0 400 20 9/22/2011 Series 2 - Wing - Fuselage x 0 5 0 330 21 9/22/2012 Series 2 - Wing - Fuselage x 0 10 0 330 22 9/22/2013 Series 2 - Wing - Fuselage x 5 0 0 330 23 9/22/2014 Series 2 - Wing - Fuselage x 10 0 0 330 24 9/22/2015 Series 2 - Wing - Fuselage x 20 0 0 330 25 9/22/2016 Series 2 - Wing - Fuselage x -10 0 0 330 26 9/23/2008 Series 2 - Wing - Fuselage x 0 0 0 250 27 9/23/2008 Series 2 - Wing - Fuselage x 0 0 0 330 28 9/23/2008 Series 2 - Wing - Fuselage x 0 0 0 400 29 9/23/2008 Series 2 - Wing - Fuselage x 0 -5 0 330 30 9/23/2008 Series 2 - Wing - Fuselage x 0 -10 0 330 31 9/23/2008 Series 2 - Wing - Fuselage x 0 5 0 330 32 9/23/2008 Series 2 - Wing - Fuselage x 0 10 0 330 33 9/23/2008 Series 2 - Wing - Fuselage x 0 0 -10 330 34 9/23/2008 Series 2 - Wing - Fuselage x 0 -5 -10 330 35 9/23/2008 Series 2 - Wing - Fuselage x 0 -10 -10 330 Continued 69 Table A.2 Continued
36 9/23/2008 Series 2 - Wing - Fuselage x 0 5 -10 330 37 9/23/2008 Series 2 - Wing - Fuselage x 0 10 -10 330 38 9/23/2008 Series 2 - Wing - Fuselage x -10 0 0 330 39 9/23/2008 Series 2 - Wing - Fuselage x 5 0 0 330 40 9/23/2008 Series 2 - Wing - Fuselage x 10 0 0 330 41 9/23/2008 Series 2 - Wing - Fuselage x 20 0 0 330 42 9/24/2008 Series 2 - Wing - Fuselage x 0 -5 -10 330 43 9/24/2008 Series 2 - Wing - Fuselage x 0 -10 -10 330 44 9/24/2008 Series 2 - Wing - Fuselage x 0 5 -10 330 45 9/24/2008 Series 2 - Wing - Fuselage x 0 10 -10 330 46 9/25/2008 Series 2 - Wing - Fuselage x 0 0 0 300 47 9/25/2008 Series 2 - Wing - Fuselage x 0 0 0 400 48 9/25/2008 Series 2 - Wing - Fuselage x 0 0 0 500 49 9/25/2008 Series 2 - Wing - Fuselage x 0 5 0 330 50 9/25/2008 Series 2 - Wing - Fuselage x 0 10 0 330 51 9/25/2008 Series 2 - Wing - Fuselage x -10 0 0 330 52 9/25/2008 Series 2 - Wing - Fuselage x 5 0 0 330 53 9/25/2008 Series 2 - Wing - Fuselage x 10 0 0 330 54 9/25/2008 Series 2 - Wing - Fuselage x 20 0 0 330
70
APPENDIX B
EXPERIMENTAL RESULTS
71 Table B1: Experimental Results Series 1 – Effects of Billow
CD AOA 1.00 1.025 1.05 1.10 0.08 0.0227 0.0155 0.0134 0.0208 2.43 0.0261 0.0176 0.0205 0.0462 4.82 0.0284 0.0201 0.0205 0.0644 7.20 0.0315 0.0200 0.0191 0.0998 9.51 0.0396 0.0258 0.0221 0.1360 11.76 0.0549 0.0346 0.0256 0.1557 13.97 0.0743 0.0501 0.0376 0.1504 16.10 0.1005 0.0787 0.0494 18.08 0.1264 0.0994 0.0659 20.08 0.1492 0.1261 0.1320 22.07 0.1759 0.1265 0.1276 24.05 0.1939 0.1279 0.1386 26.03 0.1900 0.1496 0.1271 28.01 0.2125 0.1770 0.1446 29.98 0.2280 0.2066 0.1630 31.99 0.2518 0.2159 0.2069
CL AOA 0.08 -0.0213 -0.0225 -0.0238 -0.0184 2.43 0.0429 0.0049 -0.0270 0.0067 4.82 0.1013 0.0500 -0.0150 0.1587 7.20 0.1784 0.0957 0.0161 0.3104 9.51 0.2530 0.1624 0.0662 0.3968 11.76 0.3418 0.2417 0.1365 0.3446 13.97 0.4167 0.3090 0.2084 0.2593 16.10 0.4965 0.4255 0.2817 18.08 0.5673 0.4822 0.3646 20.08 0.5890 0.5613 0.5921 22.07 0.6183 0.4969 0.5075 24.05 0.6145 0.4399 0.4932 26.03 0.5538 0.4693 0.4208 28.01 0.5556 0.4729 0.4442 29.98 0.5444 0.4970 0.4432 31.99 0.5524 0.4673 0.4848 L D AOA 0.08 -0.9387 -1.4515 -1.7725 -0.8844 2.43 1.6454 0.2790 -1.3145 0.1444 4.82 3.5722 2.4923 -0.7299 2.4661 7.20 5.6580 4.7860 0.8417 3.1092 9.51 6.3873 6.3054 2.9945 2.9184 11.76 6.2253 6.9844 5.3325 2.2135 13.97 5.6084 6.1701 5.5419 1.7245 Continued
72 Table B.1 Continued
16.10 4.9415 5.4093 5.7011 18.08 4.4897 4.8509 5.5336 20.08 3.9473 4.4514 4.4863 22.07 3.5144 3.9281 3.9777 24.05 3.1689 3.4390 3.5582 26.03 2.9153 3.1363 3.3099 28.01 2.6139 2.6714 3.0728 29.98 2.3878 2.4058 2.7185 31.99 2.1937 2.1647 2.3428
Table B2: Experimental Results Series 1 – Effects of Reducing Wing Deformation
CD AOA 0 Ribs 1 Ribs 3 Ribs 7 ribs 0.08 0.0227 0.0259 0.0220 0.0210 2.43 0.0261 0.0245 0.0139 0.0232 4.82 0.0284 0.0276 0.0166 0.0287 7.20 0.0315 0.0416 0.0324 0.0426 9.51 0.0396 0.0523 0.0467 0.0678 11.76 0.0549 0.0792 0.0770 0.1060 13.97 0.0743 0.1084 0.1192 0.1584 16.10 0.1005 0.1576 0.1695 0.2079 18.08 0.1264 0.2221 0.2251 0.2647 20.08 0.1492 0.2364 0.2613 0.2977 22.07 0.1759 0.2675 0.3054 0.2902 24.05 0.1939 0.2392 0.2923 0.3278 26.03 0.1900 0.3106 0.3328 0.3828 28.01 0.2125 0.3295 0.3744 0.4106 29.98 0.2280 0.3502 0.3975 0.4514
CL AOA 0 Ribs 1 Ribs 3 Ribs 7 ribs 0.08 -0.0213 -0.0778 -0.0836 -0.0200 2.43 0.0429 0.0449 0.1038 0.1532 4.82 0.1013 0.1959 0.2881 0.3477 7.20 0.1784 0.3236 0.4436 0.5468 9.51 0.2530 0.5131 0.6595 0.7300 11.76 0.3418 0.6459 0.8259 0.9141 13.97 0.4167 0.8130 0.9918 1.0236 16.10 0.4965 0.9785 1.1117 1.1194 18.08 0.5673 1.0462 1.1450 1.1722 20.08 0.5890 0.9532 1.1694 1.1517 22.07 0.6183 0.9555 1.1452 0.9893 24.05 0.6145 0.7860 1.0092 0.9955 26.03 0.5538 0.8879 1.0009 1.0143 28.01 0.5556 0.8936 1.0336 0.9980 29.98 0.5444 0.8381 1.0005 1.0166 Continued
73 Table B.2 Continued
L D AOA 0 Ribs 1 Ribs 3 Ribs 7 ribs 0.08 -0.9387 -3.0076 -3.7954 -0.9513 2.43 1.6454 1.8316 7.4530 6.6120 4.82 3.5722 7.1075 17.3741 12.1108 7.20 5.6580 7.7863 13.6778 12.8325 9.51 6.3873 9.8024 14.1270 10.7704 11.76 6.2253 8.1594 10.7236 8.6199 13.97 5.6084 7.5034 8.3228 6.4617 16.10 4.9415 6.2085 6.5595 5.3835 18.08 4.4897 4.7116 5.0864 4.4277 20.08 3.9473 4.0324 4.4746 3.8683 22.07 3.5144 3.5715 3.7499 3.4095 24.05 3.1689 3.2862 3.4528 3.0365 26.03 2.9153 2.8581 3.0075 2.6495 28.01 2.6139 2.7120 2.7605 2.4306 29.98 2.3878 2.3932 2.5170 2.2522
Table B.3: Experimental Results Series 2 – Effects of Leading Edge Sweep
15
2 CAOA CL CD CM C L 1.11 -0.2292 0.0125 -18.3092 -0.0535 0.0525 3.11 -0.0528 0.0102 -5.1620 -0.0946 0.0028 5.11 0.2444 0.0217 11.2520 -0.1204 0.0597 7.11 0.5748 0.0514 11.1760 -0.1573 0.3304 9.11 0.8668 0.0974 8.9019 -0.1929 0.7513 11.11 1.1475 0.1574 7.2899 -0.2095 1.3167 13.11 1.3826 0.2258 6.1229 -0.2568 1.9115 15.11 1.5099 0.2924 5.1634 -0.2772 2.2798 17.11 1.5274 0.3488 4.3784 -0.3079 19.11 1.4388 0.3769 3.8176 -0.3721 21.11 1.1856 0.3387 3.5007 -0.4461 30
2 CAOA CM C L 0.15 -0.2092 0.0152 -13.7565 -0.0230 0.0438 2.56 -0.0322 0.0135 -2.3884 -0.0532 0.0010 5.04 0.1934 0.0235 8.2205 -0.0792 0.0374 7.44 0.4136 0.0439 9.4259 -0.1017 0.1711 9.79 0.6673 0.0718 9.2991 -0.1434 0.4453 12.09 0.7795 0.1084 7.1914 -0.1737 0.6076 Continued 74 Table B.3 Continued
14.31 0.9462 0.1531 6.1823 -0.2227 0.8953 16.41 1.0983 0.2084 5.2699 -0.2714 1.2062 18.39 1.2333 0.2748 4.4878 -0.2958 20.41 1.3008 0.3396 3.8301 -0.3175 22.24 1.3072 0.3863 3.3835 -0.3826 24.35 1.2048 0.3994 3.0168 -0.4351 26.28 1.1435 0.4126 2.7715 -0.4832 45
L 2 CAOA CL CD D CM C L 1.68 -0.2446 0.0191 -12.7956 -0.0920 0.0598 3.68 -0.1350 0.0131 -10.3014 -0.0998 0.0182 5.68 0.0214 0.0138 1.5470 -0.0944 0.0005 7.68 0.1851 0.0222 8.3459 -0.1050 0.0343 9.68 0.3287 0.0377 8.7131 -0.1254 0.1080 11.68 0.4697 0.0597 7.8678 -0.1401 0.2206 13.68 0.5326 0.0716 7.4359 -0.1934 0.2837 15.68 0.6330 0.1026 6.1711 -0.2255 0.4007 17.68 0.7532 0.1447 5.2042 -0.2507 19.68 0.8474 0.1918 4.4178 -0.2736 21.68 0.9799 0.2672 3.6680 -0.2907 23.68 1.0783 0.3369 3.2005 -0.3162 25.68 1.1116 0.3848 2.8890 -0.3864 27.68 1.0510 0.4046 2.5972 -0.4144 29.68 1.0117 0.4278 2.3649 -0.4370 60
1.11 -0.1989 0.0255 -7.8015 0.0020 0.0395 3.11 -0.0760 0.0224 -3.3963 -0.0416 0.0058 5.11 0.0848 0.0242 3.4968 -0.0774 0.0072 7.11 0.2630 0.0369 7.1282 -0.0917 0.0692 9.11 0.4150 0.0577 7.1936 -0.0961 0.1722 11.11 0.5124 0.0807 6.3528 -0.1508 0.2625 13.11 0.5823 0.1017 5.7265 -0.2052 0.3390 15.11 0.6422 0.1267 5.0689 -0.2424 0.4124 17.11 0.7087 0.1575 4.4991 -0.3404 19.11 0.7717 0.1940 3.9775 -0.3890 21.11 0.9982 0.2949 3.3849 -0.3975 23.11 1.0514 0.3433 3.0624 -0.4677 25.11 1.2128 0.4431 2.7372 -0.4969 Continued
75 Table B.3 Continued
27.11 1.3899 0.5602 2.4809 -0.5546 29.11 1.3861 0.6169 2.2468 -0.5481
Table B.4: Experimental Results Series 2 – Effects of Canard Deflection - Elevator
CM - 15 CAOA e 10 e 0 e 5 e 10 e 20 1.11 -0.0609 -0.0535 -0.0432 -0.0345 -0.0404 3.11 -0.1022 -0.0946 -0.0813 -0.0718 -0.0734 5.11 -0.1299 -0.1204 -0.1104 -0.0991 -0.1030 7.11 -0.1605 -0.1573 -0.1363 -0.1252 -0.1280 9.11 -0.1892 -0.1929 -0.1788 -0.1493 -0.1547 11.11 -0.2173 -0.2095 -0.2142 -0.1806 -0.2069 13.11 -0.2506 -0.2568 -0.2508 -0.2128 -0.2419 15.11 -0.2552 -0.2772 -0.2818 -0.2413 -0.2784 17.11 -0.2959 -0.3079 -0.3137 -0.2926 -0.3210 19.11 -0.3470 -0.3721 -0.3570 -0.3445 -0.3387 21.11 -0.3878 -0.4461 -0.4108 -0.3722 -0.3997 - 30
0.15 -0.0419 -0.0413 -0.0247 -0.0070 -0.0015 2.56 -0.0732 -0.0584 -0.0418 -0.0352 -0.0049 5.04 -0.1001 -0.0626 -0.0712 -0.0634 -0.0660 7.44 -0.1265 -0.0621 -0.1104 -0.0997 -0.0984 9.79 -0.1641 -0.0836 -0.1500 -0.1439 -0.1479 12.09 -0.1983 -0.0977 -0.1948 -0.1639 -0.1867 14.31 -0.2272 -0.1310 -0.2269 -0.2143 -0.2275 16.41 -0.2601 -0.1639 -0.2634 -0.2384 -0.2822 18.39 -0.2949 -0.1747 -0.3007 -0.2788 -0.3179 20.41 -0.3177 -0.1878 -0.3482 -0.3332 -0.3710 22.24 -0.3758 -0.2488 -0.3818 -0.3691 -0.4076 24.35 -0.3964 -0.3060 -0.4301 -0.4480 -0.4407 26.28 -0.4486 -0.3662 -0.4610 -0.4722 -0.4968 - 45
1.68 -0.1295 -0.0920 -0.0012 -0.0024 0.0003 3.68 -0.1317 -0.0998 -0.0066 0.0220 0.0108 5.68 -0.1123 -0.0944 -0.0350 0.0009 -0.0108 Continued
76 Table B.4 Continued
7.68 -0.1559 -0.1050 -0.0564 -0.0196 -0.0444 9.68 -0.2314 -0.1254 -0.0889 -0.0481 -0.0354 11.68 -0.1974 -0.1401 -0.1399 -0.0437 -0.0384 13.68 -0.2275 -0.1934 -0.1826 -0.0692 -0.0604 15.68 -0.1488 -0.2255 -0.2145 -0.1152 -0.1014 17.68 -0.1472 -0.2507 -0.2948 -0.1108 -0.1369 19.68 -0.1295 -0.2736 -0.3309 -0.1304 -0.1775 21.68 -0.1403 -0.2907 -0.3767 -0.1235 -0.1684 23.68 -0.1605 -0.3162 -0.3843 -0.0907 -0.1557 25.68 -0.2426 -0.3864 -0.4175 -0.1516 -0.2281 27.68 -0.3161 -0.4144 -0.4793 -0.2241 -0.3060 29.68 -0.3182 -0.4370 -0.5146 -0.2967 -0.3334
CM - 60 CAOA e 10 e 0 e 5 e 10 e 20 1.11 -0.0212 0.0020 0.0388 0.1403 0.0952 3.11 -0.0472 -0.0416 0.0673 0.1174 0.0601 5.11 -0.0891 -0.0774 0.0657 0.0992 0.0247 7.11 -0.1102 -0.0917 0.0721 0.0619 0.0432 9.11 -0.1066 -0.0961 0.0539 0.0512 0.0379 11.11 -0.1365 -0.1508 0.0168 0.0525 0.0256 13.11 -0.2004 -0.2052 -0.0697 -0.0282 -0.0126 15.11 -0.2447 -0.2424 -0.0967 -0.0604 -0.0891 17.11 -0.2841 -0.3404 -0.1283 -0.1387 -0.1428 19.11 -0.3270 -0.3890 -0.1555 -0.1567 -0.1884 21.11 -0.3597 -0.3975 -0.2170 -0.1086 -0.2202 23.11 -0.4048 -0.4677 -0.1301 -0.1031 -0.2283 25.11 -0.4337 -0.4969 -0.1122 -0.1246 -0.2882 27.11 -0.4401 -0.5546 -0.1379 -0.0746 -0.2013 29.11 -0.4107 -0.5481 -0.1341 -0.0737 -0.1637
Table B.5: Experimental Results Series 2 – Effects of Canard Deflection – Aileron
Cn
a 15 30 45 60 -10 x -0.1313 -0.2316 0.0000 0 0.0203 0.0134 0.0524 0.0049 5 0.0659 0.0134 0.1230 0.0188 Continued
77 Table B.5 Continued
10 0.1259 0.1640 0.2431 0.0985 20 x 0.1559 x x
Cl
a 15 30 45 60 0 -0.0037 0.0001 -0.0745 0.0003 10 -0.0015 -0.0009 -0.0387 0.0062
CY
0 0.0079 0.0069 0.0246 0.0078 10 0.0555 0.0658 0.1044 0.0710
Table B.6: Experimental Results Series 2 – Effects of Canard Deflection – Sideslip
Cn
-10 -0.4943 -0.2208 -0.2916 0 0.0134 0.0524 0.0049
-10 0.0001 -0.0745 0.0003 0 0.0028 0.0044 0.0094
-10 0.0069 0.0246 0.0078 0 -0.2350 -0.1054 -0.2153
78