The Pennsylvania State University
The Graduate School
Department of Aerospace Engineering
AERODYNAMIC EXPERIMENTS ON A DUCTED FAN IN HOVER AND
EDGEWISE FLIGHT
A Thesis in
Aerospace Engineering
by
Leighton Montgomery Myers
2009 Leighton Montgomery Myers
Submitted in Partial Fulfillment of the Requirements for the Degree of
Master of Science
December 2009
The thesis of Leighton Montgomery Myers was reviewed and approved* by the following:
Dennis K. McLaughlin Professor of Aerospace Engineering Thesis Advisor
Joseph F. Horn Associate Professor of Aerospace Engineering
Michael Krane Research Associate PSU Applied Research Laboratory
George A. Lesieutre Professor of Aerospace Engineering Head of the Department of Aerospace Engineering
*Signatures are on file in the Graduate School
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ABSTRACT
Ducted fans and ducted rotors have been integrated into a wide range of aerospace vehicles, including manned and unmanned systems. Ducted fans offer many potential advantages, the most important of which is an ability to operate safely in confined spaces.
There is also the potential for lower environmental noise and increased safety in shipboard operations (due to the shrouded blades). However, ducted lift fans in edgewise forward flight are extremely complicated devices and are not well understood.
Future development of air vehicles that use ducted fans for lift (and some portion of forward propulsion) is currently handicapped by some fundamental aerodynamic issues. These issues influence the thrust performance, the unsteadiness leading to vehicle instabilities and control, and aerodynamically generated noise. Less than optimum performance in any of these areas can result in the vehicle using the ducted fan remaining a research idea instead of one in active service.
The Penn State Department of Aerospace Engineering initiated an experimental program two years ago to study the aerodynamics of ducted lift fans. The focus of this program from its initiation was to study a single lift fan subject to an edgewise mean flow. Of particular concern was the transitional flow regime from hover to a relatively high forward speed in which a major portion of the vehicle lift is produced by the aerodynamic forces on the body. We refer to this as ducted fan edgewise flow. There are four obvious consequences of operating a ducted lift fan in edgewise (forward) flow.
First, separations off the leading portion of the duct can reduce the inflow and thus the thrust of the fan. Second, the separated flow will lead to unsteadiness which will
iii undoubtedly decrease the control authority of the vehicle. Thirdly, the outer surface of the fan shroud is likely to be fairly blunt. This body shape, together with the strong momentum drag of the lift fan outflow, produce excessive drag forces that increase the requirements of the propulsion devices. Finally, increased turbulence of the inflow will also result in increased production of aerodynamic noise.
The goals of this project are to conduct detailed experiments on several configurations of ducted lift fans in hover and edgewise flow. Single ducted lift fan configurations involve different shrouded duct shapes and rotor shapes. Rotors are tested with a range of solidities and tip clearances. Including inlet duct vents over the forward portion of the duct shroud, has the potential of reducing the problem of separated flow over the forward portion of the duct inlet, and potentially reducing the drag of the vehicle in forward flight.
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TABLE OF CONTENTS
LIST OF FIGURES ...... x
LIST OF TABLES...... xix
NOMENCLATURE ...... xx
ACKNOWLEDGEMENTS...... xxiii
Chapter 1 Introduction ...... 1
1.1 Why Ducted Fans?...... 2
1.2 Ducted Fan History...... 4
1.2.1 Fixed Wing Propulsion (Forward Flight) ...... 4
1.2.2 Rotary Wing Propulsion (Forward Flight)...... 5
1.2.3 Tilt Fan VTOL...... 6
1.2.4 Direct Lift Fan...... 6
1.2.5 Lessons Learned from History...... 8
1.2.6 Future and Current Outlook...... 9
1.3 Ducted Fan Classification...... 11
1.4 Important Ducted Fan Parameters ...... 12
1.5 Technical Approach/Specific Goals ...... 15
Chapter 2 Review of Rotor Aerodynamics...... 18
2.1 Open Rotor in Hover ...... 18
2.2 Ideal Ducted Rotor in Hover...... 20
2.3 Non-Ideal Ducted Rotor in Hover ...... 25
2.3.1 Horn[7] Method...... 25
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2.3.2 Leishman[2] Method...... 28
2.4 Hover Summary...... 32
2.5 Open Rotor in Forward Flight ...... 33
2.6 Non-ideal Ducted Rotor in Forward Flight ...... 35
Chapter 3 Description of Wind Tunnel Models...... 40
3.1 Ford Fan Model ...... 40
3.2 LL1 Series Models...... 41
3.2.1 LL1 Rotor Selection...... 41
3.2.2 LL1 Motor Selection...... 43
3.2.3 LL1 Duct Design...... 45
3.3 10 Series Models...... 48
3.3.1 Model 10-1 Duct...... 49
3.3.2 Model 10-2 Duct...... 51
3.3.3 10 Series Rotor Selection...... 54
3.3.4 10 Series Motor and Electronics Selection...... 57
3.3.5 10 Series Motor Mount...... 62
3.3.6 10 Series Isolated Rotor Model...... 65
Chapter 4 Induced Velocity Experiments...... 67
4.1 Description of Experiment...... 67
4.1.1 Description of Facility...... 68
4.1.2 Instruments Used ...... 69
4.1.3 Experiment Setup...... 69
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4.1.4 LL1 Induced Velocity Procedures...... 71
4.1.5 10-1 Induced Velocity Procedures...... 73
4.2 LL1 Induced Velocity Experiment Results ...... 77
4.3 Induced Velocity of Nominal 10 inch Rotors Operating Outside the Duct....82
Chapter 5 Thrust Experiments in Hover ...... 89
5.1 Description of Experiment...... 89
5.2 Description of Facility...... 90
5.2.1 APB Balance...... 90
5.2.2 Calibrating the APB Balance...... 93
5.2.3 Hammond Balance...... 98
5.2.4 Calibrating the Hammond Balance...... 100
5.3 Data Acquisition Systems...... 104
5.3.1 APB Data Acquisition System...... 105
5.3.2 APB LabView Software Operation ...... 108
5.3.3 Hammond Data Acquisition System...... 114
5.3.4 Hammond LabView Software Operation ...... 116
5.4 Post-Processing...... 118
5.4.1 APB Post-Processing...... 119
5.4.2 Hammond Post-Processing...... 120
5.4.3 Non-Dimensionalization of Forces and Moments ...... 121
5.5 Experiment Setup...... 122
5.5.1 LL1-1 Setup...... 123
5.5.2 10-1 Setup...... 126
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5.5.3 10-2 Setup...... 126
5.6 Experiment Procedure ...... 127
5.7 Hover Thrust Experiment Results ...... 131
Chapter 6 Forward Flight Experiments (Wind Tunnel)...... 139
6.1 Description of Experiment...... 139
6.2 Description of Facility...... 140
6.2.1 APB Wind Tunnel...... 141
6.2.2 Calibrating the Test Section...... 146
6.2.3 APB Balance...... 151
6.3 Changes to the Data Acquisition System...... 158
6.3.1 Hardware Configuration ...... 158
6.3.2 Wind Tunnel LabView Software...... 163
6.4 Post-Processing...... 169
6.4.1 Non-dimensionalization of Forces and Moments...... 169
6.4.2 Tare Drag Procedure...... 171
6.4.3 Post-Processing Code...... 175
6.5 Experiment Setup...... 186
6.6 Experiment Procedure ...... 190
6.7 Forward Flight Experiment Results...... 196
Chapter 7 Remarks and Conclusions ...... 221
7.1 Summarizing Remarks...... 222
7.2 Suggestions for Future Work...... 230
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Bibliography ...... 235
Appendix A...... 237
A.1 LL1 Series Motor Power Experiment...... 237
A.2 Preliminary 10 Series Model Designs ...... 243
Appendix B...... 245
B.1 Setup of Pulleys for Calibration of APB Balance...... 245
B.2 APB Balance Calibration Curves...... 248
B.3 History of Past APB Balance Calibrations ...... 251
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LIST OF FIGURES
Figure 1-1: Ryan XV-5...... 5
Figure 1-2: Piasecki Pathfinder...... 5
Figure 1-3: Bell X-22 A...... 6
Figure 1-4: Hiller Flying Platform...... 7
Figure 1-5: Piasecki Model 59 Skycar...... 7
Figure 1-6: The X-35 using a lift fan to take off vertically...... 10
Figure 1-7: The Moller Skycar uses 4 tilt fans...... 10
Figure 1-8: X-Hawk uses 2 lift fans and 2 ducted fans for forward flight...... 11
Figure 1-9: 2007 Vertical Lift Urban Aeronautics X-Hawk...... 12
Figure 1-10: Inlet lip radius shown as r lip provided by Pereira[3]...... 13
Figure 1-11: Effect of lip radius on static thrust efficiency provided by Parlett[4].....13
Figure 1-12: Effect of tip clearance on thrust coefficient provided by Martin[5]...... 14
Figure 1-13: Effect of diffuser angle on slipstream provided by Leishman[2]...... 15
Figure 2-1: Hovering open rotor slipstream...... 18
Figure 2-2: Hovering ideal ducted rotor slipstream...... 21
Figure 2-3: Hovering non-ideal ducted rotor slipstream...... 25
Figure 2-4: Leishman’s non-ideal ducted rotor in hover...... 28
Figure 2-5: Important results of the open rotor and ideal ducted rotor in hover...... 32
Figure 2-6: Important results of the non-ideal ducted rotor in hover...... 33
Figure 2-7: Open rotor in forward (edgewise) flight...... 34
Figure 2-8: Ducted rotor in forward flight...... 36
Figure 3-1: LL1 foam-PVC interface ...... 47
Figure 3-2: Fully assembled LL1 series model...... 48
x
Figure 3-3: Model 10-1 shape transition...... 51
Figure 3-4: Forward vents on front duct of Bell Helicopter/Urban Aeronautics X- Hawk...... 52
Figure 3-5: CAD drawing of proposed model 10-2...... 52
Figure 3-6: Cut profile-view of model 10-2 ...... 53
Figure 3-7: Pitch distribution of 10 series rotors ...... 56
Figure 3-8: Thickness distribution of 10 series rotors ...... 56
Figure 3-9: Chord distribution of 10 series rotors ...... 57
Figure 3-10: Motor/ESC connection...... 59
Figure 3-11: Servo tester connection with ESC...... 60
Figure 3-12: Elogger connection with ESC...... 62
Figure 3-13: New 10 series motor mount ...... 63
Figure 3-14: Fully assembled model 10-1 ...... 64
Figure 3-15: Fully assembled model 10-2 ...... 64
Figure 3-16: Aluminum adapter for isolated rotor model...... 65
Figure 3-17: Isolated rotor model ...... 66
Figure 4-1: Three mounting positions of models...... 70
Figure 4-2: Location of mini-vane anemometer during velocity measurement of LL1 model ...... 71
Figure 4-3: Location of mini-vane anemometer during velocity measurement of 10 series model without the duct ...... 74
Figure 4-4: Induced velocity setup for 10 series isolated rotor ...... 75
Figure 4-5: Inflow ratio versus non-dimensional radius for isolated 14.5”x 11”, 4 blade rotor with tapered tips ...... 77
Figure 4-6: Inflow ratio versus non-dimensional radius for isolated 14.5”x 12”, 4 blade rotor with square tips...... 78
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Figure 4-7: Mean inflow ratio versus non-dimensional radius for both 4 blade, isolated 14.5” tapered and square tip rotors...... 79
Figure 4-8: Mean inflow ratio versus non-dimensional radius for both 4 blade, ducted 14.5” tapered and square tip rotors ...... 80
Figure 4-9: Comparison of inflow ratio for ducted and isolated 4 blade, 14.5” rotors with tapered and square tips ...... 81
Figure 4-10: Inflow ratio (at 1” above rotor) versus non-dimensional radius for the isolated 9.5” APC, 4 blade rotor...... 83
Figure 4-11: Inflow ratio (at 1” below rotor) versus non-dimensional radius for the isolated 9.5” APC, 4 blade rotor...... 84
Figure 4-12: Actual induced velocity of isolated 9.5” APC, 4 blade rotor...... 85
Figure 4-13: Comparison of inflow ratios at 1 inch and 5 inches below isolated 9.5” APC, 4 blade rotor ...... 86
Figure 4-14: Actual induced velocity of isolated 9.5” MA, 3 blade rotor...... 87
Figure 4-15: Inflow ratio (at 5” below rotor) versus non-dimensional radius for the isolated 9.5” MA, 3 blade rotor ...... 88
Figure 5-1: Force transfer schematic of APB balance...... 91
Figure 5-2: Angle of attack adjustment bar of APB balance...... 92
Figure 5-3: Pulley configuration for pure drag calibration...... 96
Figure 5-4: Drag channel calibration curves for APB balance...... 97
Figure 5-5: Hammond force balance layout, provided by Litz[15]...... 98
Figure 5-6: Pure lift calibration of Hammond balance ...... 102
Figure 5-7: Pure drag calibration of Hammond balance...... 103
Figure 5-8: Pure pitch calibration of Hammond balance...... 103
Figure 5-9: Basic DAQ structure ...... 105
Figure 5-10: DAQ setup at APB...... 106
Figure 5-11: Connector block switch configuration for hover thrust experiments...... 107
Figure 5-12: Elogger hook-up for hover thrust experiments at APB...... 108 xii
Figure 5-13: Front panel of LabView software for hover thrust experiments at APB...... 110
Figure 5-14: Front panel of LabView software for the calibration of the APB balance ...... 113
Figure 5-15: Type Caption Here...... 115
Figure 5-16: Front panel of LabView software for hover thrust experiments at Hammond ...... 117
Figure 5-17: Front panel of LabView software for the calibration of the Hammond balance ...... 118
Figure 5-18: Model LL1 setup for hover thrust experiments on Hammond balance ..124
Figure 5-19: Model LL1 setup for hover thrust experiments on APB balance ...... 125
Figure 5-20: The forward vents of model 10-2 can be opened (Left) and closed (Right)...... 127
Figure 5-21: Hover thrust coefficient versus rotor tip Mach number for 14.5” tapered tip, 4 blade rotor with and without LL1 duct...... 131
Figure 5-22: Hover thrust coefficient versus rotor tip Mach number for 14.5” square tip, 4 blade rotor with and without LL1 duct ...... 133
Figure 5-23: Hover thrust coefficient versus rotor tip Mach number for 14.5” tapered tip, 4 blade rotor with LL1 duct measured at Hammond and APB ...... 134
Figure 5-24: Tip clearance effect with model 10-1, 9.5” APC, 4 blade rotor ...... 135
Figure 5-25: Tip clearance effect with model 10-1, 9.5” MA, 3 blade rotor (Non- dimensional) ...... 136
Figure 5-26: Tip clearance effect with model 10-1, 9.5” MA, 3 blade rotor (dimensional) ...... 137
Figure 5-27: Comparison of thrust coefficient for model 10-2 with MA rotor when forward vents are open and closed...... 138
Figure 6-1: Top view schematic diagram of APB wind tunnel ...... 142
Figure 6-2: APB wind tunnel fan and stators...... 143
Figure 6-3: Measurement of turbulence intensity of APB wind tunnel, Brophy[18] ..146
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Figure 6-4: Diagram of APB wind tunnel for calibration...... 148
Figure 6-5: APB balance supported under the test section...... 152
Figure 6-6: Simulated load setup ...... 153
Figure 6-7: Initial APB lift channel simulated load results ...... 154
Figure 6-8: Preloaded weights added to the center of the APB balance...... 155
Figure 6-9: APB lift channel simulated load after added weights to center of table...156
Figure 6-10: Negative drag simulated load on APB balance...... 157
Figure 6-11: Positive pitching moment simulated load on APB balance ...... 157
Figure 6-12: BNC 2090 configuration for wind tunnel test at APB...... 159
Figure 6-13: Elogger control panel...... 160
Figure 6-14: APB wind tunnel amplifier ...... 161
Figure 6-15: APB wind tunnel connector block ...... 162
Figure 6-16: APB DAQ connection diagram for wind tunnel experiments ...... 163
Figure 6-17: APB wind tunnel LabView program front panel...... 164
Figure 6-18: Pressure difference & temperature display of APB LabView program..166
Figure 6-19: Block diagram of the APB wind tunnel LabView program: Pressure transducer calibration (Left), test section calibration constant K (Right) ...... 168
Figure 6-20: Tare drag setup for ducted fan models in APB wind tunnel...... 172
Figure 6-21: Tare drag versus wind tunnel speed for the ducted fan models in the APB wind tunnel...... 173
Figure 6-22: Tare drag setup for isolated rotor model in APB wind tunnel...... 174
Figure 6-23: Tare drag versus wind tunnel speed for the isolated rotor setup in the APB wind tunnel...... 175
Figure 6-24: Block diagram of APB wind tunnel post-processing code ...... 177
Figure 6-25: Continued block diagram of wind tunnel post-processing code...... 178
Figure 6-26: Example LabView output file from APB wind tunnel test...... 179
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Figure 6-27: Raw data input section of APB wind tunnel post-processing code ...... 180
Figure 6-28: Averaging section of the APB wind tunnel post-processing code...... 182
Figure 6-29: Quick-look graphing section of APB wind tunnel post-processing code...... 184
Figure 6-30: 10 series model electronics setup...... 187
Figure 6-31: Wire routing to angle of attack adjustment bar for 10 series ducted fan model inside APB wind tunnel...... 188
Figure 6-32: Open vent configuration of model 10-2 inside the APB wind tunnel.....189
Figure 6-33: Closed vent configuration of model 10-2 inside the APB wind tunnel ..189
Figure 6-34: Isolated rotor model mounted inside the APB wind tunnel...... 190
Figure 6-35: Wind tunnel motor calibration ...... 192
Figure 6-36: Dimensional lift as a function of wind tunnel speed for the isolated rotor, isolated duct, and ducted fan models ...... 197
Figure 6-37: Lift (non-dimensionalized with wind tunnel speed) as a function of wind tunnel speed for each ducted fan model at 6000 RPM ...... 198
Figure 6-38: Lift (non-dimensionalized with wind tunnel speed) as a function advance ratio for each ducted fan model, V WT and V Tip variable...... 200
Figure 6-39: Lift (non-dimensionalized with rotor tip speed) as a function of wind tunnel speed for each ducted fan model at 6000 RPM ...... 201
Figure 6-40: Lift (non-dimensionalized with rotor tip speed) as a function of advance ratio for each ducted fan model, V WT and V Tip variable...... 202
Figure 6-41: Dimensional lift and drag as a function of wind tunnel speed for each ducted fan model at 6000 RPM ...... 203
Figure 6-42: Comparison of dimensional drag for each isolated duct at 0 o angle of attack...... 204
Figure 6-43: Comparison of dimensional drag for isolated rotor and each ducted fan model at 6000 RPM...... 205
Figure 6-44: Drag coefficient, corrected for tare drag, as a function of wind tunnel speed for each ducted fan model at 6000 RPM ...... 206
xv
Figure 6-45: Drag coefficient as a function of advance ratio for each ducted fan model, V WT and V Tip variable...... 207
Figure 6-46: Dimensional pitching moment about 3/4 chord of each isolated duct at 0 o angle of attack as a function of wind tunnel speed...... 208
Figure 6-47: Dimensional pitching moment for each ducted fan model and isolated rotor at 6000 RPM as a function of wind tunnel...... 209
Figure 6-48: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) for each ducted fan model at 6000 RPM as a function of wind tunnel speed ...... 210
Figure 6-49: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) for each ducted fan model as a function of advance ratio, V WT and VTip variable...... 211
Figure 6-50: Pitching moment coefficient (non-dimensionalized with rotor tip speed) for each ducted fan model at 6000 RPM as a function of wind tunnel speed ...... 212
Figure 6-51: Pitching moment coefficient (non-dimensionalized with rotor tip speed) for each ducted fan model as a function of advance ratio, V WT and VTip variable...... 213
Figure 6-52: Drag coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable ...... 214
Figure 6-53: Side force coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable ...... 215
Figure 6-54: Lift coefficient (non-dimensionalized with wind tunnel speed) of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable ...... 216
Figure 6-55: Lift coefficient (non-dimensionalized with rotor tip speed) of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and VTip variable...... 216
Figure 6-56: Rolling moment coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable...... 217
Figure 6-57: Yawing moment coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable...... 218
xvi
Figure 6-58: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) of 10-1 ducted fan as a function of advance ratio and angle of attack, VWT and V Tip variable ...... 220
Figure 6-59: Pitching moment coefficient (non-dimensionalized with rotor tip speed) of 10-1 ducted fan as a function of advance ratio and angle of attack, VWT and V Tip variable ...... 220
Figure 7-1: Exit control vane design for single fan model ...... 231
Figure 7-2: Fully assembled dual fan model...... 232
Figure 7-3: Isolated dual rotor model and incidence angle adjustment...... 232
Figure 7-4: Comparison of measured power spectra for the 3-bladed MA ducted and isolated rotors...... 233
Figure 7-5: Disk loading comparison of several ducted fans ...... 234
Figure A-1: Block diagram of power curve experiment with LL1 14.5” rotor with tapered tips...... 237
Figure A-2: Power versus RPM for 14.5” x 11” rotor with a 1:1 motor gear ratio ....239
Figure A-3: Power versus RPM for 14.5” x 11” rotor with a 1:1 motor gear ratio ....241
Figure A-4: Temperature of motor with isolated 14.5” x 11” propeller and a 3.8:1 gear ratio for a duration of 20 minutes ...... 242
Figure A-5: Alternate design of 10 series ducted fan models (baseline)...... 243
Figure A-6: Alternate design of 10 series ducted fan models (change in inlet lip radius) ...... 244
Figure A-7: Alternate design of 10 series ducted fan models (increased diffuser angle) ...... 244
Figure B-1: Positive pure drag calibration setup ...... 245
Figure B-2: Positive pure side force calibration setup...... 245
Figure B-3: Negative pure lift calibration setup ...... 246
Figure B-4: Positive pure roll calibration setup...... 246
Figure B-5: Positive pure pitch calibration setup ...... 247
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Figure B-6: Positive pure yaw calibration setup...... 247
Figure B-7: APB calibration curves for pure drag loading...... 248
Figure B-8: APB calibration curves for pure side force loading ...... 248
Figure B-9: APB calibration curves for pure lift loading ...... 249
Figure B-10: APB calibration curves for pure roll loading...... 249
Figure B-11: APB calibration curves for pure pitch loading...... 250
Figure B-12: APB calibration curves for pure yaw loading ...... 250
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LIST OF TABLES
Table 3-1: Geometric parameters of Ford Fan Model ...... 40
Table 3-2: Rotors Used in Ducted Fan Research...... 42
Table 3-3: Motor Trade Study ...... 44
Table 3-4: Duct Geometries Used in Other Ducted Fan Research ...... 46
Table 3-5: LL1 Geometric Parameters ...... 47
Table 3-6: Parameters of 10 Series Rotors ...... 55
Table 4-1: Induced Velocity Experiment Test Matrix for LL1 model ...... 72
Table 4-2: Induced Velocity Experiment Test Matrix for the 10 Series Isolated Rotor ...... 76
Table 5-1: Test Matrix for Hover Thrust Experiments with model LL1...... 129
Table 5-2: Test Matrix for Hover Thrust Experiments with model 10-1 ...... 130
Table 5-3: Test Matrix for Hover Thrust Experiments with model 10-2 ...... 130
Table 6-1: Model 10-1 forward flight test matrix...... 194
Table 6-2: Model 10-2 forward flight test matrix...... 195
Table 6-3: Isolated rotor forward flight test matrix ...... 196
Table A-1: Extrapolated power required to achieve 10000 RPM with 14.5” rotor and 1:1 gear ratio ...... 239
Table A-2: Extrapolated power required to achieve 10000 RPM with 14.5” rotor and 3.8:1 gear ratio ...... 241
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NOMENCLATURE a Local speed of sound in air, γRT cD Duct chord
Dm CD Drag coefficient not corrected for tare drag, 2 5.0 ρVWT S DF
S Drag coefficient corrected for tare drag, C − C * S (C D)corrected D ()D tare S DF
Dtare (C D)tare Tare drag coefficient, 2 5.0 VWT S S
L (C L)Tip Lift coefficient, 2 5.0 ρVWT S DF
L (C L)WT Lift coefficient, 2 5.0 ρVTip S DF
l (C l)WT Rolling moment coefficient, 2 5.0 ρVWT SDF RD
m (C m)Tip Pitching moment coefficient, 2 5.0 ρVTip S DF RD
m (C m)WT Pitching moment coefficient, 2 5.0 ρVWT S DF RD
n (C n)WT Yawing moment coefficient, 2 5.0 ρVWT S DF cD
S (C S)WT Side force coefficient, 2 5.0 ρVTip S DF
xx
L CT Thrust coefficient in hover, 2 5.0 ρVTip S DF
D Duct diameter at rotor plane
Dm Force measured by horizontal load cell of balance
Dtare Measured tare drag force l Moment measured by roll load cell of balance
L Force measured by vertical load cell of balance
VTip MTip Rotor tip Mach number, a n Moment measured by yaw load cell of balance
R Universal gas constant (air)
RD Duct radius
RR Rotor radius
RPM R Rotor RPM
2 SDA Ducted fan disk area, πRD
SDF Ducted fan external wetted area, πDcD
SS Wetted area of model support structure t Duct thickness
T Local ambient air temperature vi Rotor induced velocity
xxi
RPM VTip Rotor tip speed, 2π R R 60 D
VWT Wind tunnel speed
vi λ Inflow ratio, VTip
ρ Local ambient air density
VWT Advance ratio, VTip
xxii
ACKNOWLEDGEMENTS
This project was sponsored by the Office of Naval Research, with Technical
Monitor Mr. John Kinzer. Dr. Judah Milgram and Dr. Naipei (Peter) Bi from the Sea
Based Aviation Division at the Naval Surface Warfare Center, Carderock are
acknowledged for their guidance and assistance.
I would like to thank all the students of the Ducted Fan Team at Penn State for
their support throughout the project: Kateryna Karachun, Michael Mcerlean, Lee Gorny,
Kyle Bachstein, Ben Davis, Patrick DeAngelis, Vince Dutcavich, Phil Sibley, Sohn
Ilyoup, Kim Seung Pil, Liam Brett-Eiger, Nathan Depenbusch, Jason Chauvin, Ryan
Hook, Nate Morgan, Greg Davis, Nick Hoburn, Russell Powers, and Ryan Stanley. The
efforts of these students have been invaluable to the progress of this project and it has been a very rewarding experience working with them.
I would also like to thank Mr. Wook Rhee and Mr. Richard Auhl for all their
advice, especially in the laboratory. I would like to acknowledge my advisor, Dr. Dennis
K. McLaughlin. I am grateful to him for giving me the opportunity to lead this project.
His knowledge, imparted to me, of project planning and “how to see the whole picture” is
something I will take with me long after my Penn State career.
Finally, I would like to thank my family and friends for their loving support. You
will always inspire me to achieve my dreams and beyond.
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Chapter 1
Introduction
In a world where vertical take off and landing, VTOL, missions are critical, it
appears that helicopters have become the primary vehicle of choice. Throughout the
history of such vehicles however, there has been a number of alternative solutions to the
vertical flight problem. One such alternative is known as the ducted fan. The
development of the earliest ducted fan vehicles dates to the 1940s at a time when
helicopters were reaching successful milestones. It was the hope that ducted fans would become a useful counterpart, but as the helicopter continued to gain ground, ducted fan projects were stalled and eventually abandoned. Few successes and little research have been conducted since the 1970s; however, there has been a recent resurgence of ducted
fan work in industry and academia. If the ducted fan is to be included in more
applications in aerospace vehicles, some research and development work in specific problem areas will be required.
This thesis reports on a series of experiments on model ducted lift fans. These
fans (or rotors) are models of components envisioned to be the major lift fans for future
VTOL aeronautical vehicles. The experiments were designed to explore problem areas
with ducted lift fans that perhaps have been barriers to their more widespread use in
current vehicles. The problem areas addressed in this thesis are confined to those related
to aerodynamics, including performance and parameters that relate to control. The
1 experimental program has included the design and fabrication of models suitable for wind tunnel and hover experiments.
This introduction includes brief descriptions of aerospace vehicles that have and continue to use ducted fans for both vehicle lift and forward thrust. The ducted fans described herein will be divided into categories and the specific application this research will focus on, namely vehicle lift fans, will be highlighted. Following these descriptions, the specific goals of this research will be stated along with an explanation of the technical approach.
Chapter 2 will summarize the basic principles and analysis of ducted fan/rotor
aerodynamics. Chapter 3 will provide a description of the facilities including the
description of the wind tunnel models. Chapters 4, 5, and 6 describe the various
experiments: the fan induced velocity measurements, the hover experiments, and the
wind tunnel test. Finally Chapter 7 includes discussion of the results, the conclusions
drawn and suggestions for future work. For the remainder of Chapter 1, some background information on ducted fan vehicles follows.
1.1 Why Ducted Fans?
To begin with, the mere definition of ducted fan has been confusing as to when to
distinguish between ducted fans, ducted propellers, shrouded rotors, shrouded propellers
and so on. In reality, there is no distinction between these various names. Hovey[1]
describes the device as a, “mechanically driven, single stage unit, with a central propeller
having any number of blades, surrounded by a close fitting shroud ring.” The general
2 function of the ducted fan is to produce a force by accelerating air through a ducted propeller and expelling the air downstream at the exit of the duct. Presented in this section are some of the advantages of using ducted fans.
As will be pointed out in more detail later, the duct around the propeller actually produces thrust augmentation. Hence, a smaller ducted rotor can be used and still achieve the same amount of thrust as a larger free propeller. This is beneficial for scenarios when a lot of obstacles exist in the air space. For example a small, compact aircraft would be ideal for crowded urban environments.
The idea of shrouding a rotating propeller not only protects against blade strikes with objects at low altitude, such as power lines, but it also protects people around the aircraft. Helicopters have a serious disadvantage in degraded visual environments due partly to the large main and tail rotors being exposed. Shipboard operations are especially dangerous to the personnel working around the aircraft. The ducted propeller aircraft could reduce these problems. Also in the event of blade failure, and ultimate separation, the shroud can act as a shield to any object in the area.
The mechanical complexity of helicopter hubs creates performance penalties due to drag and weight. A ducted fan aircraft may eliminate the use of hinges and bearings for control. Instead, cascades of airfoils can be added to the inlet and/or exit of the ducted fan in order to vector the flow. In doing so, control of aircraft pitching, rolling, yawing, and side force can be achieved. The reduced complexity may also mean less maintenance costs.
It is also thought that by shrouding a propeller, some acoustic shielding will be observed. Although a high number of blades produces a high frequency noise, this noise
3 is more readily subject to atmospheric attenuation with propagation distance.
Leishman[2] describes a method being used to reduce ducted fan noise by employing unequal blade spacing of the propeller.
1.2 Ducted Fan History
Upon reviewing the history and use of ducted fans for aircraft, Hovey[1] suggests
that there were four distinct categories for the type of aircraft where they were applied.
The use of the ducted fan really depended on what the particular vehicle design was. The
four categories defined by Hovey[1] are: fixed wing propulsion, rotary wing propulsion,
tilt fan VTOL, and direct lift fans.
1.2.1 Fixed Wing Propulsion (Forward Flight)
The 1 st aircraft to fly successfully through the use of a ducted fan was the
Caproni-Campini CC-1 in the year 1940. In the 1950s, the Aerodyne aircraft utilized a
circular ducted wing with exit control vanes to assist in vertical takeoffs. Ducted fan-in-
wing concepts, such as the Vanguard Omniplane 2C and the Ryan XV-5A, were also
considered as a means to provide vertical takeoff capability for fixed wing aircraft. The
Ryan XV-5A, pictured in Figure 1-1, specifically suffered from several instabilities in
hover and transition from hover to forward flight.
4
Figure 1-1: Ryan XV-5.
1.2.2 Rotary Wing Propulsion (Forward Flight)
Ducted fans have also found use in rotary wing applications. In the early 1960s, a
ducted fan was used for thrust compounding the Piasecki Pathfinder as shown in Figure
1-2. Another aircraft to utilize this type of ducted fan was the Canadian Avian Aircraft
Ltd. Model 1/180 gyroplane. Ducted fans were also incorporated into tail rotors to
counteract main rotor torque as early as 1971 with the Aerospatialle Gazelle.
Figure 1-2: Piasecki Pathfinder.
5
1.2.3 Tilt Fan VTOL
The tilt fan concept was first introduced in 1956 with the Doak VZ-4. Two other
tilt fan vehicles were operated in the 1960s including the Nord 500 Cadet and the Bell X-
22 A, (Figure 1-3). As shown in Figure 1-3, this type of aircraft featured a set of ducted
fans that were positioned upright for a vertical takeoff and then rotated horizontal to provide thrust in forward flight.
Figure 1-3: Bell X-22 A.
1.2.4 Direct Lift Fan
In the late 1950s, several aeronautical vehicles were proposed which used ducted
fans for providing lift. The Hiller Flying Platform was one of these concepts. As shown
in Figure 1-4, the pilot stood directly over the rotating propellers. In order to maneuver
in forward flight, the pilot was required to lean to a side in order to shift the vehicle’s
center of gravity.
Other ducted lift fan vehicles included the Chrysler VZ-6 and the Piasecki air jeeps. These aircraft featured dual ducted fans in a tandem configuration with the 6
Figure 1-4: Hiller Flying Platform. pilot seated in the center of the vehicle. The Piasecki Model 59 Skycar, shown in Figure
1-5, achieved lateral motion through longitudinal control vanes. Pitch control of the aircraft was obtained through differential collective rotors and lateral control vanes.
Since there was no separate thrust source, the entire aircraft was required to tilt in order to maintain forward flight. Piasecki addressed this by canting the aft fan in the successor to the Model 59, the AirGeep II. This improved the forward flight characteristics and also allowed the AirGeep II to be maneuvered more easily while on the ground.
Figure 1-5: Piasecki Model 59 Skycar.
7
1.2.5 Lessons Learned from History
(1) Poor Engine Performance
It was shown that in 1940, the CC-1 simply could not compete with other aircraft
of the time due to the unavailability of an engine to efficiently propel the aircraft. Even
engine upgrades were made to the Piasecki Pathfinder to improve forward flight speed.
The Flying Hiller Platforms met their demise partly due to low power and Piasecki added
another engine upgrade to the AirGeep II.
(2) Poor Stability During Transition Flight
This problem certainly came into play with the Ryan XV-5A/B and the Doak VZ-
4 tilt fan. Without the use of a stability augmentation system to correct frequently abrupt
and random oscillations, the pilot was left with a very uncontrollable aircraft. Transition
from hover to forward flight and back has historically been a source of aircraft instability
due to a number of factors from low forward flight speed to duct drag.
(3) Hot Gas Ingestion
This was a common problem with vehicles hovering low to the ground as hot
exhaust would recirculate back into the engine and severely degrade performance. This
was very evident with the Ryan XV-5A/B. Also sometimes the hot exhaust gases would be strong enough to erode the runway surface.
(4) Low Control Power from Vanes
Although several different configurations and arrangements were used
successfully, the control vane system was sometimes not powerful enough to do what
they intended. For example, the Ryan XV-5A/B had poor short take offs because the
8 maximum deflection angle of the vanes was only 45 o and their rotation rate was too slow to produce enough positive lift to shorten the distance as compared to a conventional takeoff.
(5) Induced Forces and Drag Due to Ducts
The Doak VZ-4 and the Bell X-22A especially had problems achieving high flight speeds do the large ducts producing too much drag. Also the two vehicles were restricted to operation in little to no crosswind. A moderate sized wind gust put such a side force on the large ducts, that it was quite difficult to control.
(6) High Fuel Consumption
If the disk area of the ducted propeller is reduced to a certain value below the required disk area of a free propeller to produce the same thrust, the power requirements will increase. A notorious criticism of the Piasecki air jeeps was that they consumed large amounts of fuel.
1.2.6 Future and Current Outlook
A new era of ducted fan use has taken advantage of past design ideas and has
improved upon them. Presented next is a few examples of the latest aerospace vehicles
that include the use of ducted fans.
In 2001, the Lockheed Martin X-35, as seen in Figure 1-6, completed its first
vertical take off. The X-35 uses a lift fan located in the fuselage just aft of the cockpit to
assist with vertical takeoffs.
9
Figure 1-6: The X-35 using a lift fan to take off vertically.
In 1997, the 12 passenger Eurocopter EC-155 was first flown using a ducted fan
for the tail rotor. In 2001, Eurocopter also released its EC-130. Both the EC-155 and
EC-130 boast lower vibrations and noise with the ducted fan tail rotor.
The final two future ducted fan vehicles are currently under development. Figure
1-7 shows the Moller Skycar. Powered by ethanol fuel, the Moller Skycar is to be
marketed as an affordable personal transportation vehicle. The ducted fans are upright
for vertical takeoff and then rotate to a forward flight position.
Figure 1-7: The Moller Skycar uses 4 tilt fans.
10
The X-Hawk, as shown in Figure 1-8, is under development by Urban
Aeronautics. The X-hawk is being marketed as a multi-purpose vehicle capable of air
rescue, pursuit, and scout missions. The two ducted fans located in the fore and aft positions of the fuselage provide lift. The use of control vanes at the duct inlet and exit
allow the vehicle to maneuver in any position. Forward flight is achieved by two rear
mounted ducted fans. By enclosing the rotors, the X-hawk can operate in areas that
would normally be inaccessible to helicopters.
Figure 1-8: X-Hawk uses 2 lift fans and 2 ducted fans for forward flight.
1.3 Ducted Fan Classification
As it has been shown, there have been many different vehicle designs. However,
all ducted rotor air vehicles can ultimately be categorized into two groups or classes.
There are those whose main source of forward thrust as well as lift comes from the
ducted rotor. For example, the Piasecki Skycar would fall into this group. There are also
those vehicles whose main source of forward thrust comes from a propulsive device other
than the ducted rotor. The ducted rotor in this case is mainly producing vertical lift. The
11
Urban Aeronautics X-Hawk would be included in this category. The aerodynamics of both of these types of aircraft is distinctly different.
The focus will now be turned to evaluating air vehicles which utilize ducted rotors primarily for lift as illustrated by Figure 1-9. This means that in forward flight, the
ducted rotor is in edgewise flow.
Figure 1-9: 2007 Vertical Lift Urban Aeronautics X-Hawk.
1.4 Important Ducted Fan Parameters
In order to construct a general framework for the design of ducted fans, this
section introduces several specific parameters. The inlet lip radius is defined as the
distance from the center of the duct inlet to the inlet edge as shown in Figure 1-10.
12
Figure 1-10: Inlet lip radius shown as r lip provided by Pereira[3].
Hovey[1] explains that the inlet lip radius should be large enough to maximize static
thrust while keeping in mind the increasing drag with size. A good range of lip radius is
from 5% to 15% of the propeller diameter. Figure 1-11 shows how the static thrust
efficiency varies with lip radius.
Figure 1-11: Effect of lip radius on static thrust efficiency provided by Parlett[4].
13
Figure 1-11 shows that as the inlet lip diameter increases, the static thrust efficiency
increases, (Parlett[4]).
Tip clearance describes the gap between the propeller blade tips and the inner
wall of the duct. The goal is to ensure that the tip gap is sufficiently small so as to
decrease the vortex structures at the blade tips and increase thrust. Hovey[1] suggests
that the blade tip clearance should be no greater than 0.015 inches for fans up to 18
inches in diameter and no greater than 0.03 inches for a 60 inch fan. Figure 1-12 shows
the effect of blade tip clearance on the coefficient of thrust as measured by Martin[5].
Seen in Figure 1-12, as the tip clearance becomes smaller, the coefficient of thrust
increases.
Figure 1-12: Effect of tip clearance on thrust coefficient provided by Martin[5].
The slipstream shown in Figure 1-13 for the ducted propeller case can be altered by adjusting the angle of the duct exit. This angle is referred to as the diffuser angle.
14
Instead of the slipstream remaining constant at the duct exit, it may actually expand with
increasing diffuser angle. This is also shown schematically in Figure 1-13.
Figure 1-13: Effect of diffuser angle on slipstream provided by Leishman[2].
Special care should be taken however to ensure that the diffuser angle does not become
so great that flow separation occurs at the duct exit.
Control vanes provide a method for vectoring the flow coming into or out of a
ducted fan. Therefore, the control vanes can be either mounted at the duct inflow or exit plane. In many cases, the vanes are located at the duct exit to provide a means of
controlling pitch, roll, or yaw moments.
1.5 Technical Approach/Specific Goals
In this research, a number of geometric arrangements or components of ducted
fan vehicles will be subjected to aerodynamic analysis through experiments (either in a
15 wind tunnel or on a hover test stand). In most cases all of these studies will be conducted over a wide range of flow fields including hover, edgewise forward flight, and transition from hover to forward flight. Measurements will consist of forces and moments. The initial concentration will be on lift, drag, and pitching moment, the latter being most critical to the stability of ducted fan vehicles. Parameters to be varied in the experiments will include rotor RPM and model angle of attack – pitch to oncoming flow.
Because of the fundamental nature of the project, the focus of the experiments will be with “generic” single ducted fans ranging in diameter of approximately 10 inches to 15 inches. This size fan can be tested in the Penn State wind tunnel (whose test section is 3.7 by 4.2 ft) and produce results which will be reasonably accurate at high forward speeds (advance ratios). It is expected that the hover experiments on the hover test stand will produce the most accurate experimental data, while the hover experiments within the wind tunnel test section will be less accurate (in representing a full size vehicle, not constrained by wind tunnel, or other walls).
Other studies will include the mean velocity profiles of the flow fields around the fan measured over a substantial inflow ratio range, simulating hover flight. These measurements will be made with a mini-vane anemometer.
The response of the fan to such parameter variation, in addition to the ratio of free stream to fan tip speeds will provide strong evidence of the degree of instability of any ducted fan vehicle throughout a major part of its perceived operating regime. Thus the specific goals of this research are:
16
(1) Design and fabricate single ducted fan models that will be used for hover and
wind tunnel testing. Flexibility will need to be built into the designs to facilitate
modification of critical parameters.
(2) Make local flow field measurements as close as possible to the rotor plane (above
and below it) with a mini-vane anemometer. This will demonstrate aerodynamic
effects in the fan flow.
(3) Perform hover and wind tunnel experiments with single ducted fan models. The
models will utilize interchangeable rotors, with 3 and 4 blades. These rotors have
different pitch and twist distributions and duplicate additional rotors have slightly
smaller diameters to produce different tip gap clearances. The exterior shapes of
the two ducts will be distinctly different in an effort to evaluate the effect of the
exterior shape on the aerodynamic properties.
17
Chapter 2
Review of Rotor Aerodynamics
2.1 Open Rotor in Hover
In order to understand the effects of adding a shroud around a rotor, an open rotor
case is first considered. The slip stream of a hovering rotor is shown in Figure 2-1.
T
Disk Area, A 1 2 vi
∞ w Figure 2-1: Hovering open rotor slipstream.
In Figure 2-1, station 1 is just above the rotor plane, station 2 is just below the rotor plane, and station ∞ is at the ultimate wake. T is the rotor thrust, vi is the induced
velocity of the rotor, and w is the velocity of the ultimate wake. Assuming, 1-D
incompressible, quasi-steady flow, the mass flow through the rotor can be determined
through the conservation of mass.
& m = ρA∞ w = ρA2vi = ρAvi (2-1)
18
Equation 2-1 states that the mass flow at the ultimate wake is equal to the mass flow at
the rotor. Essentially, the same mass that enters the rotor leaves the rotor.
The conservation of momentum can then be used to relate the rotor thrust to mass
flow. because the momentum flux well above the rotor is negligible.
T = m& w (2-2)
Finally, conservation of energy is used to relate thrust and induced velocity to mass flow.
1 Tv = m& w2 (2-3) i 2
If equations 2-2 and 2-3 are combined, the following relationship between the induced velocity at the rotor and the velocity at the ultimate wake can be obtained.
1 v = w (2-4) i 2
In other words, the velocity at the ultimate or far wake is twice the velocity induced by the rotor. This is a significant relationship for the open rotor case and will be used to simplify the next few relations. Knowing this, the conservation of mass can be revisited to relate the area of the slipstream at the rotor and the ultimate wake.
A 1 ∞ = (2-5) A 2
So the area of the ultimate wake slipstream is half the area of the rotor. It has already been seen by equation 2-4 that the velocity of the ultimate wake is greater than the velocity at the rotor. Hence, the air is accelerated through the rotor and into the ultimate wake. This area change is also seen in Figure 2-1.
19
Now returning to the conservation of momentum, the induced velocity of the rotor can be related to the rotor thrust.
T v = (2-6) i 2ρA
As the rotor thrust increases, the induced velocity will also increase. However, if the rotor area increases, the induced velocity will decrease. Equation 2-6 is a very important relationship also and it will later be seen how the induced velocity changes when a shroud is enclosed around a rotor. Finally the ideal induced power can be found.
T 2/3 P = Tv = (2-7) ind i 2ρA
Equation 2-7 shows the ideal induced power. Note that no viscous effects have been taken into consideration with this momentum theory analysis. Equation 2-7 will also be important in terms of comparing to the ducted rotor case.
2.2 Ideal Ducted Rotor in Hover
Using the same momentum theory analysis as in section 1, similar relationships can be derived for the ducted rotor. Although momentum theory does not change for different situations, the way it is applied to ducted fans does. This chapter will present two different momentum theory approaches to the ducted rotor in hover for comparison purposes. This first will follow Horn’s[7] approach.
20
Figure 2-2 shows the slipstream of an ideal ducted rotor in hover. Notice that the addition of the duct causes the slipstream at the exit to be constant and equal to the exit area of the duct. The flow does not contract like it does in the open rotor case.
TD
TR
Figure 2-2: Hovering ideal ducted rotor slipstream.
The momentum theory approach is used again here in order for comparison. The same assumptions that were made before are still valid along with a new assumption that includes the effect of the duct. Horn[7] states that the overall thrust is made of thrust contributions from the duct as well as the rotor.
T = TD + TR (2-8)
Also, for the ideal ducted case,
TD = TR (2-9) so that
T = 2TR (2-10)
21
Since the thrust of the duct is equal to the thrust of the rotor, then the overall thrust is equal to two times the thrust of the rotor. Now using this relationship in the conservation of momentum, the overall thrust can be related to the ultimate wake velocity.
& T = 2TR = mw (2-11)
Then the conservation of energy relates the work done on the rotor to the gain in energy of the fluid per unit time.
1 T v = m& w2 (2-12) R i 2
It’s the combination of equations 2-11 and 2-12 that lead to the first major difference between the open rotor and ducted rotor cases. When equations 2-11 and 2-12 are combined, a relationship between the induced velocity of the rotor and the ultimate wake velocity is found.
vi = w (2-13)
So in the case of an ideal ducted rotor, the induced velocity is equal to the velocity of the ultimate wake. This may have been slightly intuitive from Figure 2-2.
This relationship was then used to find the obvious result that:
A ∞ = 1 (2-14) A
In a way, it was already noted that the area of the ultimate wake section is the same as that of the rotor. This relationship shows that this is indeed the case. Continuing with the momentum theory analysis, the conservation of momentum is used again.
22
Inserting equation 2-13 into equation 2-11, the induced velocity can be found in terms of overall thrust.
T v = (2-15) i ρA
Equation 2-15 is similar to the result for the open rotor case (equation 2-6), but there is one difference. Equation 2-15 differs by a factor of 2 . This means that for the ideal ducted rotor in hover, the induced velocity will be greater than that of the open rotor by more than 40%. This is seen by the ratio of the ducted induced velocity to the open rotor induced velocity.
v iD = 2 = .1 41 (2-16) vio
Finally, the conservation of energy is used again to determine the ideal induced power of the rotor.
T 2/3 P = T v = (2-17) ind R i 4ρA
Again, comparing to the open rotor case, there is a difference. Equation 2-17 shows that for the same thrust and disk area, the ideal ducted rotor will require less power than an open rotor. This is further examined by finding the ratio of the ducted rotor induced power to the induced power of the open rotor.
P 1 iD = = .0 707 (2-18) Pio 2
Therefore it is possible for the ideal ducted rotor of the same disk area to see a
30% power savings over the open rotor to produce the same amount of thrust.
23
Furthermore, it can be seen from equation 2-17, that if the disk area of the ducted rotor is half the disk area of the open rotor, the same amount of power will be required in order to produce identical amounts of thrust.
T 2/3 T 2/3 P = = = P (2-19) D 4ρA 2/ 2ρA o
This could offer potential weight savings. Here is a tradeoff in the design of ducted rotors. A desired approach would be to select a disk area for the ducted rotor that optimizes the reduction in power and weight. However, as this section states, this is for the ideal ducted rotor. The assumption that the duct thrust contribution is equal to the rotor thrust contribution is not true due to non-uniform inflow and thrust. Therefore, next the case of the non-ideal ducted rotor is presented as laid out by Horn[7] .
24
2.3 Non-Ideal Ducted Rotor in Hover
2.3.1 Horn[7] Method
Figure 2-3 shows the slipstream of the non-ideal ducted rotor. Notice that the exit flow area is not constant as was the case for the ideal ducted rotor. In reality, the exit flow will contract to some degree.
TD
TR
Figure 2-3: Hovering non-ideal ducted rotor slipstream.
As before, there were some extra assumptions used in the momentum theory analysis to take into consideration the thrust contributions of the duct and rotor. This approach is no different, however a non-dimensional factor is now introduced to account for non-ideal inflow and thrust which is caused by the duct as stated earlier. This factor is called the thrust augmentation effect or kaug . So the thrust contribution of the duct is
now:
TD = kaugTR (2-20)
and the total thrust is,
25
T = TD + TR = kaugTR + TR = 1( + kaug )TR (2-21)
From equation 2-21, the range of kaug is seen.
0 < kaug < 1 (2-22)
For example if the value of kaug is 1, then the ideal ducted rotor case is seen as the total
thrust would be equal to two times the thrust of the rotor. On the other end, if kaug is 0, then the case of the open rotor is seen and the overall thrust is just the thrust of the rotor.
So for the non-ideal ducted rotor, the duct produces some percentage less than 100% of the thrust produced by the rotor. Now momentum theory can be used to relate the overall thrust to the velocity of the ultimate wake.
& T = 1(+ kaug )TR = mw (2-23)
Using the conservation of energy shown in equations 2-3 and 2-12, with equation
2-23, the relationship between the induced velocity and ultimate wake velocity can be found.
2v w = i (2-24) 1+ kaug
As seen in equation 2-24, the kaug term affects the calculation of the ultimate wake
velocity. It is interesting to note that if kaug is equal to 1, the ideal ducted rotor case is
seen and the ultimate wake velocity becomes equal to the induced velocity. It is also
seen that if kaug is equal to 0, the open rotor case is given and the result of the ultimate wake velocity being equal to two times the induced velocity of the rotor is upheld. This observation is made if the ratio of ultimate wake area to rotor disk area is found.
26
A 1+ k ∞ = aug (2-25) A 2
Now combining equations 2-23 and 2-24, the induced velocity can be determined as a
function of overall thrust.
T 1( + k ) v = aug (2-26) i 2ρA
Equation 2-26 shows that the induced velocity for the non-ideal ducted rotor is
still greater than the induced velocity of the open rotor, but it is less than that of the ideal
ducted rotor. Notice also that the induced velocity for both the open rotor and ideal
ducted rotor cases can be backed out of equation 2-26 if the proper limits of kaug are
selected. Finally, conservation of energy is used to find the ideal induced power of the
rotor.
T 2/3 Pind = TRvi = (2-27) 1(2 + kaug )ρA
Again, the kaug factor comes into play with the ideal induced power. Although the
actual value of kaug affects the outcome, it is clear that the required power of the non-ideal ducted rotor is still less than that of the open rotor for any thrust level. Depending on what the final value of kaug is, the non-ideal ducted rotor will act more like an open rotor,
an ideal ducted rotor, or somewhere in between the two. Thus, the thrust augmentation
term, kaug , is an effective method of modeling the non-ideal properties of a ducted rotor.
27
2.3.2 Leishman[2] Method
For comparison purposes, a second non-ideal ducted rotor model is presented as laid out by Leishman[2] . Leishman’s model uses the same assumptions of steady 1-D incompressible, irrotational flow. A momentum theory analysis is also used along with the Bernoulli equation. The main difference between the Horn and Leishman models is in the way they account for non-ideal effects. Horn[7] uses the non-dimensional thrust augmentation factor, kaug , that essentially allows the duct thrust contribution to be a
certain percentage, less than 100%, of the rotor thrust. The Leishman[2] model uses the
non-dimensional wake contraction parameter, aw, which allows the area of the ultimate
wake to equal a certain percentage of the rotor disk area. A nice feature of this model is
that it allows the ducted rotor’s performance to be calculated for not only if the exit wake
area contracts but also if the exit wake area expands. Figure 2-4 helps to illustrate the
model.
TR
w
Figure 2-4: Leishman’s non-ideal ducted rotor in hover.
28
Figure 2-4 shows a ducted rotor in hover. Station 0 represents a location far
upstream of the rotor where the flow is quiescent. Stations 1 and 2 are the locations just
above and just below the rotor respectively. Station 3 is the far wake. Leishman[2] also
states that careful consideration should be given to the design of the diffusing section of
the duct. If the expansion angle becomes too great, flow separation could occur at the
duct exit which would have a negative impact on performance.
The conservation of mass can be applied across the stations to find a relationship between the induced velocity, far wake velocity, and the wake contraction parameter.
& m = ρAvi = ρA∞ w = ρ(aw A)w (2-28)
In equation 2-28, A is the rotor disk area and A∞ is the area of the ultimate wake. Also the assumed wake contraction parameter is defined as:
A∞ = aw A (2-29) or the area of the ultimate wake is some percentage of the rotor disk area. Since the exit wake can expand or contract, the range of aw would be:
0< aw < ∞ (2-30)
Theoretically, there is no upper limit to aw, however in reality a maximum aw
would be reached after exit lip separation occurred on the duct. Rearranging equation 2-
28, a relationship between the ultimate wake velocity and induced velocity can be found.
v w = i (2-31) aw
29
In other words, the wake contraction parameter can also be defined as the ratio of the
induced velocity to the ultimate wake velocity. The conservation of momentum is then
used to relate the overall thrust to the wake contraction parameter.
ρAv 2 & i T = TD + TR = mw = (ρAvi )w = (2-32) aw
This allows us to further define the wake contraction parameter in terms of overall thrust.
ρAv 2 a = i (2-33) w T
Equation 2-33 can also be rearranged to find a definition of the induced velocity.
a T v = w (2-34) i ρA
It is seen here that if aw is equal to 1, then the induced velocity of the ideal ducted rotor is obtained as shown earlier. Also, if aw is equal to ½, the induced velocity of an open rotor is obtained. So it appears that the use of the wake contraction parameter, aw, is appropriate.
Now Bernoulli’s equation can be applied to the areas above and below the rotor
so that the thrust of the rotor can be related to aw. Equation 2-35 is obtained when
Bernoulli’s equation is applied to the area between stations 0 and 1.
1 p = p + ρv 2 (2-35) 0 1 2 i
Then Bernoulli’s equation is applied to the area between stations 2 and 3.
1 1 p + ρv 2 = p + ρw2 (2-36) 2 2 i 0 2
30
The thrust of the rotor can then be found using equations 2-35 and 2-36. This is
accomplished by taking the difference between the pressure at station 2 and station 1, and
then multiplying by the rotor disk area.
1 T = ( p − p )A = ρw2 A (2-37) R 2 1 2
If equations 2-32 and 2-37 are then compared, the ratio of rotor thrust to overall ducted
rotor thrust can be found.
T (2/1 ρAw2 ) w 1 R = = = (2-38) T ρAvi w 2vi 2aw
There are a couple of interesting points here. Again, if the correct value of a w is selected, then the ideal ducted rotor and open rotor cases can be backed out. For example, if aw is equal to 1, then the thrust of the rotor is equal to ½ the total thrust. This
would indicate that the thrust of the duct makes up the other half of the total thrust and
that the duct thrust would be equal to the rotor thrust. This is the ideal ducted rotor case.
However, if the exit wake is expanded and aw becomes greater than 1, then the thrust contribution of the rotor itself becomes a smaller percentage of the overall thrust. This would mean that the duct is contributing to the majority of the overall thrust.
Finally, equations 2-34 and 2-38 can be used to determine the induced power of the rotor.
2/3 T awT T Pind = TRvi = = (2-39) 2aw ρA 4aw ρA
31
2.4 Hover Summary
For comparison purposes as well as review, Figure 2-5 shows the key results
obtained for the open rotor and the ideal ducted rotor in hover both provided by Horn[7] .
Figure 2-5: Important results of the open rotor and ideal ducted rotor in hover.
Figure 2-6 shows the important results for the hovering non-ideal ducted rotor
obtained from Horn[7] and Leishman[2] . The two models of non-ideal ducted rotors are essentially the same. The wake contraction parameter, aw, of Leishman’s[2] method can be related to the thrust augmentation factor, kaug , of Horn’s[7] method as shown in
equation 2-40. Expansion of the duct exit could be modeled if kaug was assigned a
1( + k ) a = aug (2-40) w 2
32 value greater than 1. A value greater than 1 would also indicate that the duct produces more than twice the thrust of the rotor.
Figure 2-6: Important results of the non-ideal ducted rotor in hover.
2.5 Open Rotor in Forward Flight
The aerodynamics of ducted rotors in forward flight will now be analyzed. Again a comparison will be made with the open rotor case so that the effects of enclosing a rotor with a shroud can be seen. Therefore, a general description of open rotors in forward flight is presented.
Figure 2-7, presented by Horn[7] shows an open rotor in forward flight. As seen from Figure 2-7, the plane of the rotor is at an angle of attack, α , to the free-stream velocity, V0. When in forward flight, the rotor encounters an inflow that is spread over a
33
Figure 2-7: Open rotor in forward (edgewise) flight.
larger area than when in hover. In forward flight, a component of the free-stream
velocity also contributes to the total vertical velocity through the rotor. Thus for the same
amount of thrust, the induced component of velocity through the rotor, vi, in forward flight is less than the induced velocity of the rotor in hover. As the forward speed increases, the free-stream component will dominate the total vertical velocity through the rotor and the term vi will continue to be reduced. Recalling from equation 2-7 that,
Pind = Tvi (2-7)
it is seen that as forward speed increases and vi decreases, the induced power will decrease. This is known as the translation lift effect.
The variable χ is called the wake skew angle. This is shown in Figure 2-7 as the angle between the vertical axis of the rotor and the exit wake of the rotor and can be defined as:
−1 V0 cosα χ = tan (2-41) V0 sinα + vi
34
As the forward speed increases, the wake skew angle will go to 90 o. This has potential
importance for the ducted rotor in terms of overall control as will be discussed later. The
discussion of open rotor forward flight will conclude with a description of the steady-
state inflow equation.
Equation 2-42 shows a linear inflow distribution across the rotor. The assumption
of linear inflow is a simplification of reality, however this will provide a basis for
comparison to the ducted rotor case.
2 T 4 3 2 2 vi + 2V0 sinαvi + V0 vi − = 0 (2-42) 2ρAD
It should be noted that if the forward speed, V0, goes to zero, the induced velocity in
hover is given as seen in equation 2-6.
2.6 Non-ideal Ducted Rotor in Forward Flight
Figure 2-8 shows a ducted rotor in forward flight. As stated earlier, this type of
ducted rotor will produce its thrust from some other source external to the duct system.
Thus this ducted rotor will operate at small angles of attack since its main purpose is to produce a vertical force. With that said however, α , is still shown as the angle between
the horizontal axis and the free stream velocity in Figure 2-8 such that: r ˆ ˆ V0 = V0 cosαi −V0 sinαj (2-43)
The flow physics between the open rotor and ducted rotor are similar in edgewise
flight in that the rotor turns the flow downward; however, the two are vastly different
after the flow leaves the rotor and the duct further turns the flow to align it along the 35 duct’s vertical axis. Once the flow exits the duct, the wake skew angle will be affected by this extra turning of the flow. In order to account for this, a non-dimensional
r j T
i
Figure 2-8: Ducted rotor in forward flight.
parameter called the turning efficiency, kχ , is introduced. The range of kχ is
0 < kχ < 1 (2-44)
A value of 1 would mean that the flow has encountered 100% turning. In other words, the flow has been completely realigned parallel to the axis of the duct. In reality, the
value of kχ would never be equal to 1, but the higher it is would mean better control vane
effectiveness. A value of 0 corresponds to the case of the open rotor and control vane
effectiveness would be low since exit control vanes would not encounter any chord-wise
flow. The velocity through the ducted rotor is shown to be a function of the turning
36 efficiency in equation 2-45. Similarly, the velocity of the ultimate wake is shown to be a function of turning efficient in equation 2-46. r ˆ ˆ VR = 1( − kχ )V0 cosαi − (V0 sinα + vi ) j (2-45)
r ˆ ˆ V∞ = 1( − kχ )V0 cosαi − (V0 sinα + v∞ ) j (2-46)
Using equations 2-45 and 2-46, the wake skew angles at the exit of the duct and the ultimate wake can be found.
−1 1( − kχ )V0 cosα χ = tan (2-47) V0 sinα + vi
1( − k )V cosα −1 χ 0 χ∞ = tan (2-48) V0 sinα + v∞
Note that it is assumed that the horizontal component of velocity is the same at the duct exit and in the ultimate wake. With the addition of the turning efficiency term,
the wake skew angles are less than that of the open rotor case. Also, if kχ is equal to its
maximum value of 1, then both wake skew angles will be 0 o. This is important in terms
of control effectiveness as alluded to before. With a lower wake skew angle, the
effectiveness of control vanes in the wake will be greater. This explains one of the
reasons why control vanes are not used in open rotor situations but could be feasible for
the ducted rotor.
So far, the idea of turning efficiency seems to be fairly important. Perhaps the
most important aspect of the turning efficiency however is how it affects momentum
37 drag. In order to account for this term, the thrust vector of the ducted rotor was split into vertical and horizontal components so that: r ˆ ˆ T = Dmi + jT (2-49)
This can be seen in Figure 2-8. Equation 2-21 is used for the total thrust with the thrust augmentation factor included to account for non-ideal effects of the ducted rotor.
Returning again to momentum theory and the conservation of mass, the mass flow can be found. r & m = ρAVR (2-50)
Furthermore, the conservation of momentum can be used to relate the thrust vector to the velocity vectors at the inlet and at the ultimate wake. r r r & T = m(V∞ −V0 ) (2-51)
Now if only the horizontal components of the vectors in equation 2-51 are used, a relationship between the momentum drag and turning efficiency can be found.
& & Dm = m((1− kχ )V0 cosα −V0 cosα) = mkχV0 cosα (2-52)
It is seen from equation 2-52 that not only will the momentum drag increase with forward
speed, but it will also increase with turning efficiency. It is interesting to note that if kχ is
equal to 0, then the open rotor case is obtained and the momentum drag goes to 0.
Therefore, there is a tradeoff in the design of the ducted rotor for forward, edgewise,
flight in the selection of turning efficiency. The value of kχ should be optimized so that
38 the most amount of control vane effectiveness is gained while not being heavily penalized by the momentum drag.
Finally, the conservation of energy is applied to the system.
r r 1 r r r r − T ⋅V = m& (V ⋅V −V ⋅V ) (2-53) R R 2 ∞ ∞ 0 0
After inserting the necessary vectors into equation 2-53, an equation for the steady-state inflow can be obtained in the form given in equation 2-54.
T A× B − = 0 ρA
where A = 1( − k ) 2V 2 cos 2 α + (V sin α + v ) 2 χ 0 0 i (2-54)
2 k augV0 sinα − vi k augV0 sinα − vi 2 2 2 and B = − + − (k χ − 2k )V cos α χ 0 1 + k aug 1 + k aug
The non-linear steady-state inflow in equation 2-54 captures the non-ideal effects of the ducted rotor due to the thrust augmentation term and turning efficiency. The use of these terms also seems appropriate when a value of 0 is used for the forward speed. If V0 is
equal to 0, the induced velocity to hover for the non-ideal ducted rotor, as given in
equation 2-26, is obtained. With some rearranging of terms, the thrust augmentation
factor, kaug can be redefined in terms of forward speed, V0, and angle of attack, α .
V sinα v 2 0 +1 ∞ v v k = i i −1 aug 2 2 (2-55) 2 V0 cosα v∞ V0 sinα v∞ (k χ − 2k ) + 2 + χ v v v v i i i i
39
Chapter 3
Description of Wind Tunnel Models
3.1 Ford Fan Model
The first ducted fan wind tunnel model constructed at Penn State is highlighted in the thesis of Tilford[14] . Tilford[14] used the model to validate a newly refurbished six component force and moment balance. The single ducted fan model was essentially composed of a ten bladed electric automobile radiator fan and a circular duct made of
FOAMULAR 250 polystyrene foam insulation.
Looped galvanized sheet metal was used to provide a smooth surface for the interior of the duct surface as well as form a straight duct exit. Minor modifications were made to the previous design which allowed the fan to be separated from the duct and mounted to the force and moment balance. A list of geometric parameters for this initial ducted fan design is given in Table 3-1.
Table 3-1: Geometric parameters of Ford Fan Model
40
3.2 LL1 Series Models
Due to the inherent limited capabilities of the Ford Fan model, a new series of wind tunnel models would be created. The design process involved determining the size and geometry of the model as well as strategic planning of a shroud model that is both versatile in shape while being simple to use in future wind tunnel testing. It was desired to make the geometry such that shroud shape can be tailored with moderate ease. The capability of adding upstream and downstream vanes to test their effects in thrust vectoring was also considered. The model was also to incorporate various tip clearances.
Other considerations in the design included force balance attachment, material selections, mounting capability, power source selection, and minimum wind tunnel flow obstruction.
This series of ducted fan wind tunnel models were referred to as the “LL1 series.”
The LL1 models consist of three basic parts, the rotor, the motor, and the duct.
Each of these components will be outlined next.
3.2.1 LL1 Rotor Selection
A trade study was conducted to better understand the important parameters involved in selecting a rotor. A list of rotors used in ducted fan research of others was tabulated and is presented in Table 3-2.
An important factor in the design of a ducted rotor is the rotor tip clearance.
Ideally, the tip clearance should be 0, but this is obviously not possible as the blades need to be able to rotate inside of the duct. Hovey[1] suggests that the blade tip clearance
41
Table 3-2: Rotors Used in Ducted Fan Research
Tip Blade Rotor Chord Pitch [in] Tip Shape Clearance No. of Blades Solidity Thickness Rotor RPM Diameter [in] Length [c/D] [%R] [%c] 0.07 at hub to Martin [5] 9.6 - 9.9 square 1, 2, 4 2 0.1 2000 9500 0.09 at tip Collective = 5 0.05, 0.25, 0.3, Pereira [3] 6.3 square 3 0.06 0.12 5 2000 4000 to 40 degs 0.4, 0.8 2 0.005 to 0.06 * Two coaxial Sato [8] 9.8 9.6 1 0.006 to 0.06 0.02 9000 rotors, chord 0.003 to 0.03 inc then dec 0.12 at hub to Abrego [9] 37.8 fixed square 0.035 to 1.2 5 0.3 1800 3400 0.07 at tip 4 angle at 75% Parlett [10] 28 square 0.009 *tandem fan 0.1 0.25 station = 18 o model 2 0.07 at hub to Parlett [4] 18 square 0.007 *max chord at 0.06 6000 10500 0.03 at tip 50% span 0.1 at hub to Yoeli [11] 9.5 7 square 2 0.11 0.07 at tip
42 should be on the order of 0.1% and 0.2% of the rotor radius. However, Camci[12] used a tip clearance of approximately 1.0% R during a tip casing study of axial flow ducted fans.
Taking both references into consideration, it was decided to include varying tip clearances into the new model. This was accomplished by shaving the tips off a rotor that was originally larger in diameter than the duct. This also allowed for a rotor tip shape study to be performed. The largest tip clearance considered was 1.4% R. Further reductions in tip clearance could also be accomplished with the installment of a foam ring insert inside the duct in the plane of the rotor. Thus two rotors, each with four blades, were chosen with a nominal diameter of 14.5 inches. One rotor had tapered tips and 11 inch pitch. The other rotor started as a 15.5 inch rotor (12 inch pitch) but had its tips shaved so that they became square.
3.2.2 LL1 Motor Selection
A motor trade study was performed which compared rotor rotational speeds
obtained by others experimenting with similar ducted fan models. A table of motor
choices is presented in Table 3-3. Parlett[4] and Martin[5] used propeller diameters that
are the closest to that selected for the LL1 series being 18 inches and 10 inches
respectively. In both cases, the researchers ran their propellers at RPM ranges from 2000
to 10,000.
The result of the motor trade study also revealed that there are two types of
electric motors worth considering. These consist of brushed and brushless. Brushless
43
Table 3-3: Motor Trade Study
Max Power Continuous Max Burst Overall Overall Name Kv Voltage [V] ESC Prop Range* Shaft D Price [watts] Current Current Length D Trinity Co27 Monster Stock Pro Brushed N/A N/A N/A N/A N/A N/A N/A 3.17 mm 57 mm 36 mm $33 Monster Max Brushed Inrunner N/A N/A N/A N/A N/A N/A N/A 3.17 mm 57 mm 36 mm $50 ElectriFly Ammo Inrunner Brushless Motor 4875 222 7.4 to 11.1 20 30 A 36 A 8x4E to 10x7E 3 mm 33 mm 24 mm $50 E-Flite Park 370 Brushless Outrunner 1360 125 7.2 to 12 12 15 A (15 sec) 10 to 20 A 8x6 to 10x4.7 3.17mm 25mm 28 mm $50 E-Flite Park 480 Brushless Outrunner 910 250 7.2 to 12 20 25 A (15 sec) 20 to 35 A 10x7 to 12x6 4 mm 33 mm 35 mm $70 E-flite Power 46 Brushless Outrunner Motor 670 800 14.4 to 19.2 40 A 55 A (15 sec) 60 A 12x8 to 14x10 6 mm 55 mm 50 mm $110 E-flite Power 60 Brushless Outrunner Motor 400 1200 18.5 to 28.8 40 A 60 A (15 sec) 80 A 14x8 to 16x10 6 mm 62 mm 50 mm $130 E-flite Power 32 Brushless Outrunner Motor 770 700 12 to 16.8 42 A 60 A (15 sec) 60 A 11x7 to 14x10 5 mm 50 mm 42 mm $90 E-flite Power 25 Brushless Outrunner Motor 870 550 12 to 16.8 32 A 44 A (30 sec) 40 to 45 A 11x8 to 14x7 5 mm 54 mm 35 mm $85 E-flite Power 10 Brushless Outrunner Motor 1100 375 7.2 to 12 30 A 38 A (30 sec) 35 to 40 A 10x5 to 12x6 5 mm 43 mm 35 mm $75 Great Planes Rimfire Outrunner Brushless 850 815 7.4 to 14.8 45 A 55 A 11x8 to 18x10 4 mm 64.5 mm 35 mm $68 Great Planes Rimfire Outrunner Brushless 800 1480 11.1 to 18.5 50 A 80 A 10x5 to 14x7 5 mm 67 mm 42 mm $73
Kv = motor shaft RPM per input volt (without propeller) Max Burst Current = Current at full throttle *Prop Range is tabulated for 2 bladed props
44 motors are generally more efficient than brushed motors due to the benefit of not having brushes which can create losses through contact corrosion or arcing. Unfortunately, since brushless motors require AC phasing, the setup is a little more complicated and therefore expensive. Because of this, it was decided to use a cheaper brushed motor. The motor selected was the Monster Max Brushed inrunner motor.
3.2.3 LL1 Duct Design
While the inner diameter of the duct was set based on the diameter of the rotor
and tip clearance, a third trade study was performed to determine the profile shape of the
duct. Critical design parameters were defined and tabulated, comparing the geometries of
several ducted fan designs of others as shown in Table 3-4.
The duct inlet was made of the FOAMULAR 250 material which could easily be
machined to any desired shape. The ring structure of the duct consisted of PVC tubing.
The two pieces of the duct were connected together via a slot machined into the PVC
tube. The foam piece was simply slid into the slot of the PVC. This connection is shown
in Figure 3-1.
The last piece of the LL1 series model was the motor housing. A housing
structure needed to be designed and fabricated that would hold the motor/rotor
combination in the center of the duct. The housing needed to be designed in such a way
so that it could also be removed from the duct and a separate isolated rotor test could be performed. Hovey[1] explained that the hub diameter is not critical up to 40% of the
45
Table 3-4: Duct Geometries Used in Other Ducted Fan Research Researcher Model Flight Vanes Function t/d c/d rl/d rotor plane* t c rl d P. Martin [5] Duct 1 Edgewise no L and P 0.12 0.58 0.03 About 25% 1.15 5.77 0.29 10.00 Duct 2 Edgewise no L and P 0.11 0.58 0.02 About 25% 1.10 5.77 0.17 10.00 Pereira [3] LR06-D10-L72 Axial no L and P 0.28 0.72 0.06 About 20% 1.75 4.53 0.41 6.30 LR09-D10-L72 Axial no L and P 0.28 0.72 0.09 About 20% 1.75 4.53 0.56 6.30 LR13-D10-L72 Axial no L and P 0.00 0.72 0.13 About 20% 1.75 4.53 0.82 6.30 LR13-D10-L31 Axial no L and P 0.28 0.31 0.13 About 20% 1.75 1.95 0.82 6.30 LR13-D10-L50 Axial no L and P 0.28 0.50 0.13 About 20% 1.75 3.15 0.82 6.30 Sato [8] Duct 1 Axial no L and P 0.26 1.19 0.13 32(%) 2.56 11.71 1.23 9.84 Duct 2 Axial no L and P 0.25 0.94 0.13 13(%) 2.49 9.25 1.23 9.84 Abrego [9] Duct 1- "Baseline" Edgewise Flaps L and P 0.03 0.26 n/a About 50% 1.11 10.00 n/a 38.00 Duct 2- "Variation" Edgewise Flaps L and P 0.03 0.39 n/a About 50% 1.11 15.00 n/a 38.00 Parlett [4] Duct 1 Edgewise no L and P 0.03 0.68 0.01 50(%) 0.50 12.25 0.25 18.00 Duct 2 Edgewise no L and P 0.03 0.68 0.03 50(%) 0.50 12.25 0.50 18.00 Duct 3 Edgewise no L and P 0.03 0.68 0.04 50(%) 0.50 12.25 0.75 18.00 Duct 4 Edgewise no L and P 0.03 0.68 0.06 50(%) 0.50 12.25 1.00 18.00 Duct 5 Edgewise no L and P 0.03 0.68 0.08 50(%) 0.50 12.25 1.50 18.00 Parlett [10] Duct 1 Edgewise Yes L and P 0.14 0.32 0.07 90(%) 4.00 9.00 2.00 28.00 Yoeli [11] Model Edgewise no Lift 0.06 0.30 About 40% - - - 9.50
All units in inches *based on chord, starting at leading edge(%) t = thickness of duct c = chord of duct d = inside diameter of duct rl = lip radius L = Lift P = Propulsive
46
Figure 3-1: LL1 foam-PVC interface duct diameter. Thus, the hub structure was also designed to accommodate a gear box.
The motor is secured to the housing by set screws.
The motor housing was supported by three struts that are threaded into the outer shroud duct. Symmetric NACA airfoils were placed around the struts to provide a smoother internal duct flow. A table describing the final geometric characteristics of the
LL1 series model is shown in Table 3-5 and a fully assembled picture of the LL1 series ducted fan is shown in Figure 3-2.
Table 3-5: LL1 Geometric Parameters
Inner Duct Duct Expansion Diameter Chord Length Angle [in] [in] [C/D] [Deg] 14.5 4 0.28 0
47
Figure 3-2: Fully assembled LL1 series model
3.3 10 Series Models
The following section describes the design and fabrication of two additional
ducted fan wind tunnel models of a smaller scale than the LL1 series. The new, smaller
models needed to have at least the same capabilities as the LL1 series with the possibility
of some improvements. For example, a smaller scale model could provide larger ranges
of rotor RPM. This is due to the fact that lower diameter and pitch propellers require less
motor torque than their larger counterparts. It was also desired to use a better system for
controlling the rotor RPM. The LL1 series was limited in this capability. The design and
fabrication of the new 10 series models was done by a group of undergraduate aerospace
engineering students, supervised by the author. The first model to be created was simply
a scaled down version of the larger LL1 series model. The second 10 series model was to be of slightly different size and shape. Therefore, the only parameter that is distinctly
48 different between the two 10 series model is the duct. The motor and power systems are identical. Several preliminary designs for the second model are shown in Appendix A.
3.3.1 Model 10-1 Duct
Model 10-1 again features a two piece duct made of PVC tubing and
FOAMULAR 250 foam. The 10 in the name of the model stands for the approximate
inner diameter of the duct in inches. Initially, the duct shape was kept non-dimensionally
the same as the larger LL1 series. One of the advantages of designing a smaller ducted
fan model was that the model could be fabricated by students at the machine shop on
campus. The new 10 inch diameter was within the range of the machining capability.
This also reduced fabrication cost. First, the PVC tubing was cut to the appropriate chord
length with a band saw. The tube was then mounted to a lathe and the inside surface was
machined to maintain a constant inner diameter within a tolerance of 0.5 mm. The slot
for the foam duct inlet was also cut on the lathe. In order for the duct to have a sharp
trailing edge, the outside of the tube was stepped on the lathe 75% from the leading edge.
The PVC tube was then mounted onto a milling machine and a rotary index was used cut
several holes for the motor and force balance attachments.
The foam for the duct inlet was first cut to a square and then cut to a rough outer
diameter with a band saw. The foam block was then mounted to the lathe machine at its
center point. A cutting tool was then used to create the slot for the PVC to slide into.
While still mounted on the lathe, the duct contour was cut into the foam by stepping from
inlet to exit. Finally, the PVC tube is inserted into the foam and the two were mounted
49 on a milling machine so that the inner diameter could be cut out of the foam. The foam inlet was smoothed with sandpaper.
In an attempt to make the model more professional-looking, it was decided that
the new 10-1 model should be coated in some material. The coating material should not
erode the foam and should also be able to be painted. Several options were used on test
foam pieces to determine the optimal coating material. These included industrial spray
glue, Elmer’s wood glue, Gorilla Glue, epoxy, fiberglass, Bondo, spot putty, spray primer, and spray paint. At the conclusion of these tests, it was found that the spot putty
and spray primer when applied directly to the foam will chemically erode the material.
All of the glues and epoxies were found to provide an adequate coat, but it was difficult
to prevent dripping around the curved surface of the duct. An exception to this was the
spray glue, but it left the surface too tacky. The optimal case was found to be composite
fiberglass. This would provide extra strength for the model and also protect it during
handling. Several strips of fiberglass were cut and laid out on a wax paper sheet to be
wetted with a combination of epoxy resin and hardener. The ratio of resin to hardener
was 5:2. Once the epoxy was mixed, it was applied to the fiberglass strips. The strips
were wrapped chord-wise around the circumference of the duct. After about 30 minutes
the fiberglass began to cure and 12 hours was required for full curing.
One final adjustment was made to the duct shape of 10-1. It was decided that the
shape should be different than that of the LL1 series model. The shape transition is
shown in Figure 3-3.
Therefore, several layers of Bondo were used to fill in and smooth out large gaps
in the contour of the model. Smaller gaps were filled in with automotive spot putty. The
50
Figure 3-3: Model 10-1 shape transition model was smoothed with sandpaper of various grits and a coating of spray primer was applied. This process was repeated until the desired surface smoothness was achieved. A final coat of glossy black spray paint was then applied to the model.
3.3.2 Model 10-2 Duct
The 10-2 model would have different values of chord and thickness compared to
model 10-1 in order to study the effects of differently shaped ducts. It was also decided
for comparison purposes that model 10-2 would be sized non-dimensionally similar to the
duct used in aerodynamic experiments performed by Martin[5] . Fleming[13] also noted
that a significant amount of drag and pitching moment are introduced for a ducted fan in
forward edgewise flight. One proposed solution to substantial drag forces in forward
flight is the inclusion of vents on the forward portion of the duct as shown in Figure 3-4.
Figure 3-4 shows a conceptual (non-flying) mock-up of the Bell Helicopter-Urban
Aeronautics X-Hawk. This is a dual fan air vehicle with ducted fans in a tandem
51
Figure 3-4: Forward vents on front duct of Bell Helicopter/Urban Aeronautics X-Hawk configuration providing lift. This air vehicle is marketed for logistics and assault support.
It is believed that the inclusion of the forward vents could reduce inlet lip separation, thereby reducing asymmetric duct lift, drag, and ultimately pitching moment. Thus it was decided to incorporate forward vents into the design of model 10-2. Figure 3-5 shows a
CAD view of the proposed model 10-2.
Figure 3-5: CAD drawing of proposed model 10-2
52
The duct shape was fabricated in the same fashion as model 10-1; however, the addition of the forward vents introduced another complexity in the fabrication process.
Also since the chord of model 10-2 was larger, a two piece foam inlet had to be made.
For model 10-2, foam made up the entire exterior shape of the duct. This eliminated the excessive use of Bondo as before. The PVC ring was still used to provide a strong interior structure and the model was again wrapped in fiberglass. A cut-profile view of model 10-2 is shown in Figure 3-6.
Figure 3-6: Cut profile-view of model 10-2
In order to cut the forward vents into the model, special tooling had to be created that would hold the model securely in the milling machine. With this secure mounting, the vents were cut separately in longitudinal sweeps. A smaller drill bit was used for more precise cutting; however, due to the roundness of the bit, the corners of the vents
53 were not rectangular. When the vents were cut, the foam was left exposed. This would cause undesirable effects if the un-coated foam was painted. Therefore, micro-balloons filled with epoxy resin was used to coat the inner cut-surface of the forward vents.
Micro-balloons consist of small glass particles that when mixed with epoxy cure to a strong, smooth surface. The micro-balloon resin was applied to each side of the inner- vent surface separately until all four sides of the exposed foam were coated. It required about 12 hours between each coat to allow for curing. Final preparations were identical to those of 10-1 before painting. Model 10-2 was smoothed with various sizes of sand paper and an initial coat of Rustoleum brand spray primer was applied. After final sanding, the model was painted.
3.3.3 10 Series Rotor Selection
The rotor selection was made simpler by the smaller size of the model. More
options for rotor blade number and pitch setting are available for diameters of roughly 10
inches. Several rotors were again considered as shown in Table 3-6. Two rotors would be selected with different solidity. Both rotors were to be of reasonable pitch such that
they could be directly mounted to the motor. A smaller diameter rotor and smaller pitch
eliminated the need for a gear box as was used in the LL1 models. Tip clearance was
adjusted by shaving the tips off the rotors. This would also make the rotor tips
rectangular, that are more representative of conventional lift fan rotors. Thus, rotors of
slightly larger diameter than the duct were chosen and the tips were trimmed to obtain a
specific clearance. The two rotors that were selected were the 3-bladed Master Airscrew,
54
Table 3-6: Parameters of 10 Series Rotors Blade Rotor Chord Pitch [in] Tip Shape No. of Blades Solidity Thickness Rotor RPM Diameter [in] Length [c/D] [%c] 0.08 at hub to 57 at hub to 13 APC 9.5x7 9.5 7 tapered 2 0.06 2500 17000 0.02 at tip at tip 0.09 at hub to 53 at hub to 14 APC 11x7 9.55 7 square 2 0.1 2500 15000 0.07 at tip at tip 0.05 at hub to 68 at hub to 14 APC 11x6 9.57 6 rounded 4 0.14 2500 15000 0.06 at tip at tip 0.06 at hub to MA 10x5 10 5 square 3 0.1 2500 16500 0.04 at tip
55
MA, 10x5 rotor and the 4- bladed Advanced Precision Composite, APC, 11x6 rotor. The geometric characteristics of these rotors were measured and are shown in Figures 3-7, 3-
8, and 3-9. Both rotors were balanced using the modified prop balancer. Again, small amounts of material were removed from the heavier blades with sandpaper.
40.0 APC 11x6, 4 Blade 35.0 MA 10x5, 3 Blade 30.0 25.0 20.0 15.0 Pitch (Deg.) 10.0 5.0 0.0 0.0 0.2 0.4 0.6 0.8 1.0 Non dimensional radial position, r/R
Figure 3-7: Pitch distribution of 10 series rotors
0.80 APC 11x6, 4 Blade MA 10x6, 3 Blade 0.60
t/c 0.40
0.20
0.00 0.0 0.2 0.4 0.6 0.8 1.0 Non dimensional radial position, r/R
Figure 3-8: Thickness distribution of 10 series rotors
56
0.20
0.16
0.12 c/R 0.08
0.04 APC 11x6, 4 Blade MA 10x5, 3 Blade
0.00 0.0 0.2 0.4 0.6 0.8 1.0 Non dimensional radial position, r/R
Figure 3-9: Chord distribution of 10 series rotors
3.3.4 10 Series Motor and Electronics Selection
The motor trade study was revisited for the 10 series models. Both, model 10-1
and 10-2 would use the same motor. It became apparent quickly that due to the reduced
size of the model and rotor, brushless electric motors would now be a viable option.
Such small brushless motors are readily available and widely used in the remote control
airplane hobby arena. Because of this, many more documents are available describing
the specific propeller to motor combinations. This was not available for the brushed
electric motors that were used in the LL1 models. Furthermore, brushless electric motors
offer more precise control in terms of RPM. While extra components are needed for this
control, the brushless motor was chosen over the brushed motor.
Since a dc power supply was going to be used instead of batteries, the new motor
was initially sized according to the maximum current output of the power supply. It was
found from past experience that it is the maximum current that determines the maximum 57 rotor RPM. Therefore, a brushless electric motor was found with a maximum burst current (full-throttle current) in the range of 15 - 18 amps. The choice now became narrowed to brushless inrunner or outrunner motors. Whether or not a motor is classified an inrunner or outrunner has to do with the internal workings of the motor. If the motor shaft rotates inside of a stationary casing, then the motor is an inrunner. For an outrunner motor, the outer-casing of the motor physically rotates at the same rate as the motor shaft.
This can cause complications if a motor shroud is built to surround the motor. Sufficient clearance must be maintained so that the motor can operate properly. Typically inrunner motors operate at very high kV (RPM per input volts) ratings. Inrunner motors are well suited for low torque applications and gear boxes are normally applied. Outrunner motors are well suited for high torque applications and typically operate at low to moderate kV ratings. Outrunner motors are also the most common in remote controlled airplanes. Thus it was determined to use a brushless outrunner motor.
Other factors to consider when selecting this type of outrunner motor are the input voltage range and electronic speed control, ESC. The input voltage range of the motor is normally scaled with the maximum burst current. Lower maximum burst currents normally require lower input voltages. In this case, the voltage range of the power supply is sufficient up to 40 volts. Since motors with maximum burst currents on the order of 15 amps require input voltages of around 12 volts, the current power supply should work fine. Therefore, the motor that was chosen was the E-flite Park 370 outrunner with a kV of 1370 and maximum burst current of 15 amps. Next the method of connecting the motor to a DC power supply will be discussed.
58
The electronic speed control is tasked with many things. The ESC is primarily responsible for managing the speed of the motor. The most obvious function of the ESC is that it allows the motor, which has three wires, to be connected to a DC power supply, which has two terminals. Thus one end of the speed control connects to the motor and the other to the power supply. This connection is shown in Figure 3-10.
Figure 3-10: Motor/ESC connection
As shown in Figure 3-10, the blue wire coming from the motor connects to the white wire coming from the ESC. This is known as the “signal” wire. The red and black, positive and negative, wires are also connected respectively. If any one of these three wires is connected out of order, the motor rotation will simply be reversed.
The actual speed of the motor is manually controlled by a separate component
externally. The Astro Servo Tester was used for the throttle control. It is connected to
the receiver port on the ESC as shown in Figure 3-11. It is important that polarity is not
reversed when connecting the servo tester. The servo tester allows for manual adjustment
of the current pulse width that is supplied by the ESC to the motor from 1 milli-second to
59
2 milli-seconds. An alternative to this approach would be to use a wireless hand held radio commonly utilized by R/C pilots. This however would introduce extra cost and redundant features.
Figure 3-11: Servo tester connection with ESC
Other functions of the ESC include acting as a braking mechanism for the motor and automatically cutting-off power to the motor if sufficient voltage is not available.
This protects the power supply from being damaged. The ESC can also be programmed to reverse the rotation of the motor; although, it is simpler to just reverse any one of the three wires coming from the motor as explained above. The ESC also protects the motor from getting power spikes at initial start-up. Many speed controllers come with battery eliminator circuits, BECs. The BEC is an internal circuit built into the ESC. The BEC exists so that additional external power will not be required to operate the throttle control which connects to the receiver port of the ESC. However, the BEC only functions within certain motor input voltage ranges. Typically these voltage ranges are up to 13 volts.
Finally, the ESC is rated for a maximum amount of current. The maximum
60 current rating of the ESC must be at least big enough to withstand the maximum burst current of the motor. Because of this, the ESC is typically sized to about 1.5 times the maximum burst current of the motor. This provides extra protection for both the ESC and the power supply. In order to accommodate all of these features, the Castle Creations
Thunderbird 36 was chosen as the electronic speed controller for the 10 series models.
The final addition to the electronics setup of the 10 series models is the Eagle
Tree Systems Elogger. The Elogger is a commercial data logging product capable of recording performance and tracking data of remote controlled aircraft. Numerous sensors are available that allow for the measurement of rotor RPM, motor temperature, and airspeed. These measurements can be monitored in real-time through a USB link to a computer. The proprietary software also provides some post-processing support through tables, charts, and graphs. The software also allows for different models to be stored, so running different tests requires little reset time. The Elogger gets connected to the system between the ESC and power supply. As shown in Figure 3-12, the black and red power supply wires coming from the ESC get plugged into the ESC port on the Elogger.
Polarity is important here.
Special dean’s connectors have been used to ensure the correct polarity is maintained. Next to the ESC port on the Elogger is a port labeled BATT. This is where the system gets connected to the power supply. Again dean’s connectors were used to ensure the correct polarity. Also care was taken to use a male connector on the BATT end of the Elogger. The female connecter is applied to the end of the power supply wires. This will ensure safety if the power supply is turned on when holding the ends of the wires.
61
Figure 3-12: Elogger connection with ESC
3.3.5 10 Series Motor Mount
A similar motor mounting system used previously, was incorporated into the new
smaller model. Clear lexan material would again be used to shroud the motor. However,
it was decided that the new motor mount would be a permanent fixture in the 10 series
models. A separate isolated rotor model would be fabricated to determine the performance of the rotor outside of the duct. This was done to ease the design and fabrication of the new motor mount. This would also reduce setup time required during experiments.
Two iterations of the motor mounts were incorporated into the 10 series models.
The difference between the two motor mounts is in the manner that the motor physically attaches to the lexan material. The support structure for the isolated rotor model is also different from that of the ducted models. Both motor mounts feature a cylindrical shroud with an inner diameter slightly larger than the motor diameter. This would allow for the
62 outrunner motor to function without interference. The top portion of both motor mounts was also beveled to reduce sharp corners from being introduced in the downwash of the rotor. The outside shape of the new motor mount is shown in Figure 3-13.
Figure 3-13: New 10 series motor mount
For both the 10-1 and 10-2 ducted models, the motor attaches to an intermediate aluminum disk piece through four M3 0.5x20 screws. The aluminum disk has three 1/4-
20 tapped holes around its perimeter that line up with three holes drilled into the lexan motor shroud. Three threaded bolts connect the motor to the ducts of 10-1 and 10-2.
Care was taken to ensure that the motor shaft is directly in the center of the duct.
Aerodynamic streamlined tubes were again applied over the threaded bolts. With these final pieces in place, the fully assembled models 10-1 and 10-2 are shown in Figures 3-14 and 3-15 respectively.
63
Figure 3-14: Fully assembled model 10-1
Figure 3-15: Fully assembled model 10-2
64
3.3.6 10 Series Isolated Rotor Model
Since there was no duct to support the motor in the isolated rotor model, the motor shroud was modified. Instead of boring a constant diameter hole in the center of the lexan shroud as before, the inner diameter of the motor shroud was stepped. Four holes were drilled and tapped into a ledge on the inside of the lexan shroud. The motor mount supplied with the motor was then attached directly to the ledge on the inside of the lexan shroud. The motor could then be attached to the motor mount. Two threaded rods were run straight through the lexan motor shroud to hold the isolated rotor model during an experiment. An adapter, shown in Figure 3-16, made of aluminum was also fabricated to allow the isolated rotor model to be tested in three geometric positions on a test stand.
The isolated rotor model is shown in Figure 3-17.
Figure 3-16: Aluminum adapter for isolated rotor model
65
Figure 3-17: Isolated rotor model
66
Chapter 4
Induced Velocity Experiments
The theory laid out in chapter 2 shows that the addition of the shroud around the
rotor should increase the induced velocity over that of an isolated rotor. The theory also
shows that the downwash velocity, or the velocity far from the rotor plane, should
decrease when the shroud is placed around the rotor. Recall that this is because the duct
controls the shape of the slipstream after exiting the rotor. Thus this chapter highlights
velocity profile measurements of the ducted fan models, with and without the duct
attached, in a hover configuration only.
4.1 Description of Experiment
The purpose of these experiments was to quantify the velocity profile around
ducted and “un-ducted,” or isolated, rotors. The data generated from these studies could be used to validate numerical tools as well as give some insight to the physical phenomena at hand. Velocities can be sampled above and below the rotor to estimate the
induced velocity. The velocity can also be measured further downstream to estimate the
downwash velocity. The span-wise velocity distribution of the rotor can be determined by measuring the velocity at several points along the rotor.
Three different models of varying size and geometry were used in these
experiments. The velocity was measured around both the isolated rotor and ducted rotor
67 for the LL1 model. Two 14.5 inch rotors of varying pitch distribution and taper were used for the LL1 model. The velocity profile was also measured for the isolated rotor 10-
1 model. Again two different rotors were used for comparison. In this experiment three variables were introduced such that the two rotors differed by blade pitch distribution, blade taper, and number of blades . Figures 3-7, 3-8, and 3-9 show a comparison of the
10-1 rotors. For all these models, the blade tip clearance was kept fixed at 1% of the
rotor radius.
4.1.1 Description of Facility
A proper location to perform these isolated and ducted rotor tests was essential.
The safety of the experimenters was paramount. Since the rotors would be operating at
high rotational speeds, precautions needed to be taken. It was also important to protect
any experimental equipment that could be damaged during the experiment. For these
reasons, two protection screens were designed and fabricated. The screens were made of
chain-link gates each roughly 4’ by 5’. A layer of 1/8” plexiglass was also applied to the
side of each gate. Simple legs constructed of aluminum and PVC pipe were fastened to
the screens. The height of the screens can be adjusted from 5 ft. to 8 ft. The screens
enclose the ducted fan model during the experiment. It was ideal to use the chain-link
and plexiglass combination so that the model could still be viewed through the screen
during the experiment. This combination also ensured that any particle that may separate
during an experiment will not penetrate the safety screens.
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4.1.2 Instruments Used
During the experiment, the rotational rate of the fan was measured using an optical
infrared sensor made by Monarch Sensors. Reflective tape placed at the root of each blade acted as a triggering device for the optical sensor. As each piece of tape passed by
the sensor, a voltage pulse output. The frequency of this voltage pulse was then
measured by an HP spectrum analyzer. This frequency is called the blade passage
frequency, BPF. The BPF is easily converted into rotor RPM by equation 4-1. This
RPM was checked using a second infrared sensor and a variable frequency stroboscope.
BPF RPM = 60 nb (4-1)
where nb is the number of blades.
Velocity measurements were made with a propeller-type anemometer. It is called
the Testo 416 mini-vane anemometer. The anemometer features a small diameter propeller of 0.5 inches. The measurement range of the mini-vane anemometer is 0.6 to
40 m/s with a resolution of 0.1 m/s. A digital display is connected to the anemometer probe by a 3 ft. chord. The anemometer probe consists of a 1 ft. telescoping rod capable
of extending to 2 ft. maximum.
4.1.3 Experiment Setup
A special hover test stand was developed for this experiment. A solid base platform was made of oak 2 x 6 wood. The red struts from the six component force and
moment balance were affixed to the wood platform. The model then attached to the red
69 struts. This would allow for a similar test setup between the thrust measurement and induced velocity experiments.
Attachments were made for each model that allowed three different geometric positions. Figure 4-1 defines the positions. Each model can be tested in the 0 o or upright,
180 o upside down, and 90 o axial positions. These 0 o and 180 o positions were chosen to determine the effect of a ground plane. As will be discussed in the hover force measurement chapter, the 90 o position was chosen as a means to measure thrust in the
drag channel of the force and moment balance. Standard lab clamps were used to hold
the anemometer and optical sensors to stands placed around the model.
Figure 4-1: Three mounting positions of models
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4.1.4 LL1 Induced Velocity Procedures
The induced velocity was measured for the LL1 ducted fan by placing the mini-
vane anemometer at the duct exit. This was done because the anemometer could not be placed inside the duct. Therefore, the velocity was measured about 4 inches below the
rotor plane. When the induced velocity of the isolated rotor was measured, the
anemometer was placed 1 inch below the rotor plane. Figure 4-2 illustrates this setup.
Figure 4-2: Location of mini-vane anemometer during velocity measurement of LL1 model
Although there was a difference in the measurement plane, the advantage of using
the mini-vane anemometer was in the telescoping probe. The span-wise location along
the blades could be more precisely controlled with the mini-vane anemometer.
Therefore, velocity measurements were taken at 1 mm increments from blade tip to root
resulting in 14 span-wise locations. In all cases, the velocity was measured with the LL1
models mounted upside down, or exhausting to the ceiling. Since the LL1 models feature
a brushed electric motor, the RPM was controlled by stepping the input. For the LL1
71 induced velocity measurements, the input voltage was stepped from 4.0 V to 7.0 V in 1 volt increments. The input voltage was then stepped back from 7.0 V to 4.0V in 1 volt increments. This resulted in seven rotor RPM cases. An average was then taken between the two measurements at the same rotor RPM. The general procedure was to set the mini-vane anemometer at the desired span-wise location and then vary the rotor RPM.
Table 4-1 shows the test matrix for the induced velocity measurements of model LL1.
Table 4-1: Induced Velocity Experiment Test Matrix for LL1 model Duct ON/OFF Rotor Power Supply Voltage Propeller Anemometer [V] Position 4 0.3 1 5 0.3 1 14.5"x11" 6 0.3 1 4 blade 7 0.3 1 tapered tip 6 0.3 1 5 0.3 1 4 0.3 1 ON 4 0.3 1 5 0.3 1 15.5"x12" 6 0.3 1 4 blade 7 0.3 1 square tip 6 0.3 1 5 0.3 1 4 0.3 1
4 0.3 1 5 0.3 1 14.5"x11" 6 0.3 1 4 blade 7 0.3 1 tapered tip 6 0.3 1 5 0.3 1 4 0.3 1 OFF 4 0.3 1 5 0.3 1 15.5"x12" 6 0.3 1 4 blade 7 0.3 1 square tip 6 0.3 1 5 0.3 1 4 0.3 1 1. In all configurations, the rotors were positioned such that the downwash was blowing upward 2. The mini-vane anemometer was moved in 1mm increments from 0.3R to the blade tip
72
4.1.5 10-1 Induced Velocity Procedures
For measuring the induced velocity of the 10 series isolated rotor, another
velocity measurement device was introduced. Two instruments were to be used
simultaneously including the Testo mini-vane anemometer and elogger airspeed sensor.
The elogger airspeed sensor uses a pitot static probe and pressure transducer to measure
the total pressure. The elogger software then converts the total pressure to a velocity
which is monitored through a USB connection to a laptop computer. Initially the two
velocity measurement devices were positioned normal to the flow, 1 inch below the rotor
and aligned with marks drawn on the rotor at various span-wise locations. Then both
devices were positioned in a similar fashion at 5 inches below the rotor. Finally, the
mini-vane anemometer was positioned 1 inch above the rotor at each span-wise location
while the pitot-static probe of the elogger sensor was left 1 inch below the rotor. Figure
4-3 illustrates these placements.
Five locations were selected along the spans of both rotors for measuring the
induced velocity. The locations were again noted as a percentage of the rotor radius, 0 being the center hole of the rotor, and 100 being the tip of the rotor. Therefore, velocity
measurements were taken at 95%, 85%, 75%, 55%, and 33%. 33% was chosen because
it was the inner-most location that could be measured without the measurement devices
hitting the hub structure.
Since the 10 series models feature a brushless electric motor, the Astro servo
tester was used to precisely control the rotor RPM. The rotor RPM was varied from 3000
to 6000 in increments of 500 RPM. This kind of RPM control could not be obtained with
73
Figure 4-3: Location of mini-vane anemometer during velocity measurement of 10 series model without the duct the previous models. The rotor RPM was obtained by two devices. The Monarch optical sensor was used as before, but a new elogger optical sensor was also used. The elogger optical RPM sensor required placing white-out marker on each of the blades on the
“below” side of the rotor. Also, the elogger optical RPM sensor had to be placed a distance of 1 mm away from the white-out surface for optimal measurement. Again, the elogger software was used to monitor and record the rotor RPM. After inputting some parameters such as number of white-out marks, the elogger software displays a numeric value of rotor RPM.
74
The final addition to the induced velocity experiments included a temperature sensor. A small thermocouple was fastened to a non-rotating part of the motor and was connected to the elogger. Figures 4-4 shows the experimental setup of the 10 series induced velocity test.
Figure 4-4: Induced velocity setup for 10 series isolated rotor
75
Thus the velocity was measured for the 10 series isolated rotors at three positions
(1 inch below, 5 inches below, and 1 inch above) and five span-wise locations along the blade. In all cases, the isolated rotor was positioned upside down so that it was
exhausting to the ceiling. The general procedure included operating the rotor through a
range of RPMs after positioning the mini-vane anemometer and pitot-static probe at a
single span-wise location. Then the power was shut off and the anemometer and probe
were moved to the next span-wise location. The rotor RPMs were then varied again and
the process was repeated until all configurations were completed. Table 4-2 contains the
test matrix for the 10 series isolated rotors during the induced velocity experiment. Note
in Table 4-2 that some of the RPM values were skipped for certain rotors. This was because a strong vibration was noticed at an RPM of 4500 for the 4 bladed APC rotor.
Table 4-2: Induced Velocity Experiment Test Matrix for the 10 Series Isolated Rotor Isolated Rotor C Isolated Rotor D APC 11x6, 4 blade MA 10x5, 3 blade square tip square tip Anemometer/ 0.33 0.55 0.75 0.85 0.95 0.33 0.55 0.75 0.85 0.95 Pitot Probe Position, r/R* Rotor RPM 3000 4000
4500
5000
6000
* 1: The mini-vane anemometer and pitot probe were located 1 inch below rotor plane * 2: The mini-vane anemometer was located 5 inches above the rotor plane and the pitot probe was located 1 inch below the rotor plane * 3: The mini-vane anemometer and pitot probe were located 5 inches below rotor plane
** In all configurations, the rotors were positioned such that the downwash was blowing upward
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4.2 LL1 Induced Velocity Experiment Results
The measured induced velocities were non-dimensionalized with rotor tip speed, using the inflow ratio defined as:
v λ = i (4-2) vTip
Figure 4-5 shows the inflow ratio plotted against non-dimensional radius for the 14.5” x
11” isolated rotor, with tapered tips. Thus it is seen from that when the induced velocities are non-dimensionalized with rotor tip speed, the data for all corresponding rotor RPMs collapses within a close tolerance.
0.250 *Tapered Tip (11" pitch) Rotor
0.200
0.150
0.100 1740 rpm 2070 rpm 0.050 2340 rpm 2640 rpm Non dimensional induced velocity, λ Mean Value 0.000 0.00 0.20 0.40 0.60 0.80 1.00
Non dimensional radial location, r/R R
Figure 4-5: Inflow ratio versus non-dimensional radius for isolated 14.5”x 11”, 4 blade rotor with tapered tips
77
Figure 4-5 also shows that there is a positive value of velocity just outboard of the rotor hub at a span of 30%. The induced velocity steadily increases to a maximum value at around the 70% span-wise location. After this point, the induced velocity appears to decrease sharply out to the tip. Recall that this rotor has four blades with tapered tips.
Next, the same plot is presented for a four blade rotor with square tips.
Figure 4-6 shows the inflow ratio versus non-dimensional radius for the 14.5” x
12” isolated rotor, with square tips. Again, the same trend is seen in that all rotor RPMs
collapse to a single curve when the induced velocity is non-dimensionalized with rotor tip
speed. The velocity also starts positive and reaches a maximum value around the 70%
span-wise location until the velocity sharply drops off to zero at the blade tips.
0.250 *Square Tip (12" pitch) Rotor
0.200
0.150
0.100 1740 rpm 2070 rpm 0.050 2340 rpm 2550 rpm
Non dimensional induced velocity, λ Mean Value 0.000 0.00 0.20 0.40 0.60 0.80 1.00
Non dimensional radial location, r/R R
Figure 4-6: Inflow ratio versus non-dimensional radius for isolated 14.5”x 12”, 4 blade rotor with square tips
78
In order to quantitatively compare the induced velocity of each rotor, the mean inflow ratio values were superimposed over each other on the same graph.
Figure 4-7 presents the mean inflow ratio versus non-dimensional radius for each four bladed rotor. Figure 4-7 shows that the induced velocities are the same for both rotors up until a span-wise location of about 60%. Two notes of interest occur after the
60% span-wise location. First, the mean induced velocity is greater for one rotor over the other outboard of the 60% span-wise location. This can be attributed to the increase in pitch of the 14.5” x 12”, square tipped rotor. Second, the span-wise location where the maximum induced velocity occurs is shifted further outboard for the square tipped rotor.
It appears that a tapered tip tends to cause the rotor to “un-load” further inboard.
0.250 * Mean Value of each Rot. Speed
0.200
0.150
0.100
0.050 Tapered Tip (11" pitch)
Non dimensional induced velocity, λ Square Tip (12" pitch) 0.000 0.00 0.20 0.40 0.60 0.80 1.00 Non dimensional radial location, r/R R
Figure 4-7: Mean inflow ratio versus non-dimensional radius for both 4 blade, isolated 14.5” tapered and square tip rotors
79
Next the same procedure was followed for the case of the ducted rotor. Figure 4-
8 shows the mean inflow ratio versus non-dimensional radius for the two ducted rotors.
Figure 4-8 shows similar trends to the isolated rotor case. Again the increased pitch of the 14.5” x 12” rotor causes the induced velocity to increase over the entire span of the rotor. Also, its rectangular tips shift the point of maximum induced velocity further outboard. There are also some interesting differences between the isolated and ducted
0.250 * Mean Value of each Rot. Speed
0.200
0.150
0.100
Tapered Tip (11" pitch), with Duct 0.050 Square Tip (12" pitch), with Duct Non dimensional induced velocity, λ
0.000 0.00 0.20 0.40 0.60 0.80 1.00 Non dimensional radial location, r/R R
Figure 4-8: Mean inflow ratio versus non-dimensional radius for both 4 blade, ducted 14.5” tapered and square tip rotors cases. Notice in Figure 4-8 that the location of the maximum induced velocity is around the 60% span-wise location instead of 70% for the isolated rotors. This could be due to the wall of the duct at the blade tips. The wall effectively prevents the tip vortices of each blade from dissipating away from the rotor plane. The vortices appear to reflect off
80 the wall and travel back down the blade to the rotor. This can also be seen in the more docile increase in the induced velocity along the span. There is a larger area of maximum induced velocity where as the velocity reaches a distinct maximum point in the case of the isolated rotor.
Lastly, the isolated and ducted rotor cases were directly compared. Figure 4-9 shows the mean inflow ratio versus non-dimensional radius for both rotors in the ducted and un-ducted cases. Recall that ideal ducted fan theory predicts the induced velocity of the ducted rotor is larger than that of the isolated rotor. Figure 4-9 contradicts this theory.
This could be due to the unequal measurement planes. The induced velocity for
0.250 * Mean Value of each Rot. Speed
0.200
0.150
0.100
Tapered Tip (11" pitch), with Duct 0.050 Square Tip (12" pitch), with Duct Tapered Tip (11" pitch), without Duct Non dimensional induced velocity, λ Square Tip (12" pitch), without Duct 0.000 0.00 0.20 0.40 0.60 0.80 1.00 Non dimensional radial location, r/R R
Figure 4-9: Comparison of inflow ratio for ducted and isolated 4 blade, 14.5” rotors with tapered and square tips
81 the ducted case was measured at 4 inches below the rotor plane with the induced velocity for the isolated rotor case was measured at 1 inch below the rotor plane. The discrepancy could also be caused by a swirl velocity inside the duct. Wake swirl reduces the net change in fluid momentum and would have the effect of reducing the downwash velocity of the isolated fan. This is due to viscous stresses on the wall of the duct. This swirl velocity could act in such a way that the velocity measured in the plane 4 inches below the rotor would be reduced.
4.3 Induced Velocity of Nominal 10 inch Rotors Operating Outside the Duct
The induced flow of the nominal 10 inch rotors operating in hover were measured
as the first step in qualifying the 10 inch ducted fan models. Initially, the 4-bladed APC
rotor was tested following the procedure laid out earlier. In general, it was expected that
the velocity below the rotor would drop off close to the tips and that the velocity above
the rotor would not. Figure 4-10 shows the inflow ratio versus non-dimensional radius
when the mini-vane anemometer is traversed along the span at 1 inch above the rotor plane. The velocity just above the rotor plane is steady over the majority of the blade
span, but does begin to decrease toward the tip. The blades in this case featured
rectangular tips. Although, the velocity just above the rotor plane is steady, there is a point where the velocity is maximized. It appears that the point on the blade span closest
to the motor housing (around 30% r) has the highest value of velocity. This could be due
to some flow interaction with the hub structure itself. More likely however, is that the blade has a higher pitch inboard toward the rotor hub. Refer to Chapter 3 for the pitch
82 distribution of this rotor. Indeed, the maximum pitch along the blade span occurs at the
30% location.
z/R = - 0.21 APC 9.5"x 6", 4 Blade 0.25 3000 RPM 4000 RPM 0.20 5000 RPM Average 0.15
0.10
0.05 Non dimensional local velocity, λ
0.00 0.0 0.2 0.4 0.6 0.8 1.0 Non dimensional radial position, r/R R
Figure 4-10: Inflow ratio (at 1” above rotor) versus non-dimensional radius for the isolated 9.5” APC, 4 blade rotor
Figure 4-11 displays the velocity measured 1 inch below the rotor plane. Again the inflow ratio is plotted against non-dimensional radius for the 4-bladed rotor. Figure
4-11 shares the same trend as the larger rotors of the LL1 model. When the measured velocity is non-dimensionalized with rotor tip speed, the data for each RPM collapses to a single curve. Again, the point of maximum induced velocity occurs near the 75% span- wise location. After this location, the velocity begins to drop off sharply to zero at the tips.
83
z/R = 0.21 APC 9.5"x 6", 4 Blade 0.25
0.20
0.15
0.10 3000 RPM 4000 RPM
0.05 5000 RPM Average Non dimensional induced velocity, λ 0.00 0.0 0.2 0.4 0.6 0.8 1.0
Non dimensional radial position, r/R R
Figure 4-11: Inflow ratio (at 1” below rotor) versus non-dimensional radius for the isolated 9.5” APC, 4 blade rotor
The induced velocity is actually the velocity in the rotor plane. It is impossible to measure this velocity using the current techniques. The reason for measuring the velocity just above and just below the rotor is so that a better estimate of induced velocity can be
obtained. Thus, the mean inflow of both Figures 4-10 and 4-11 are plotted together
against non-dimensional radius. This is shown in Figure 4-12. The actual induced
velocity was calculated by taking the average between the velocity measured above and below the rotor plane.
84
APC 11x6, 4 Blade 0.30 Measured, z/R = - 0.21 Measured, z/R = 0.21 0.25 Actual Induced Velocity
0.20
0.15
0.10
Non dimensional local velocity, λ 0.05
0.00 0.0 0.2 0.4 0.6 0.8 1.0
Non dimensional radial position, r/R R
Figure 4-12: Actual induced velocity of isolated 9.5” APC, 4 blade rotor
Finally, the velocity was measured at 5 inches below the rotor plane for the 4- bladed rotor. This is seen in Figure 4-13. Figure 4-13 gives insight to the rotor slipstream one radius below from the rotor plane. The velocity appears to maintain a constant value from the rotor hub until just after the mid-span. The velocity then decreases steadily to zero at the tip.
The results from Figure 4-11 were also added to Figure 4-13. Thus, the data labeled “Average z/R = 0.21” represents the average inflow ratio for the 4 blade APC rotor at 1 inch below the rotor plane. This was done to show that the isolated rotor slipstream contracts with increasing distance below the rotor plane, as assumed by the momentum theory analysis of Chapter 2. It is evident that the isolated rotor slipstream
85 contracts since the location of maximum induced velocity measured at 5 inches below the rotor is further inboard from the rotor tip than that measured at 1 inch below the rotor.
z/R = 1.05 APC 9.5"x 6", 4 Blade 0.25
0.20
0.15
0.10 3000 RPM 4000 RPM 5000 RPM 0.05 Average
Non dimensional local velocity, λ Average z/R = 0.21 0.00 0.0 0.2 0.4 0.6 0.8 1.0
Non dimensional radial position, r/R R
Figure 4-13 : Comparison of inflow ratios at 1 inch and 5 inches below isolated 9.5” APC, 4 blade rotor
For comparison purpose, the same experiment was performed for the 3-bladed
MA rotor. Figure 4-14 shows the velocity measured just above and below the rotor plane as before. The actual induced velocity was calculated in the same manner as before.
Finally, the velocity was measured at 5 inches below the 3-bladed MA rotor. Figure 4-15 shows the inflow ratio versus non-dimensional radius for this location.
Both Figures 4-14 and 4-15 display the expected features of the isolated rotor test.
Of interest is the comparison of the 3-bladed rotor to the 4-bladed rotor. It is shown that the overall velocities measured for the 3-bladed rotor are less than that of the 4-bladed
86 rotor. This could be due to the extra blade. However, it has already been seen that an increase in blade pitch has a significant affect on the velocity profile. The 4-bladed rotor does indeed have a higher blade pitch distribution than the 3-bladed rotor.
MA 10x5, 3 Blade 0.30 Measured, z/R = - 0.21 Measured, z/R = 0.21 0.25 Actual Induced Velocity
0.20
0.15
0.10
0.05 Non dimensional local velocity, λ
0.00 0.0 0.2 0.4 0.6 0.8 1.0
Non dimensional radial position, r/R R
Figure 4-14: Actual induced velocity of isolated 9.5” MA, 3 blade rotor
87
z/R = 1.08 MA 9.5"x 5", 3 Blade 0.25
0.20
0.15
0.10 3000 RPM 4000 RPM 5000 RPM 0.05 6000 RPM
Non dimensional local velocity, λ Average
0.00 0.0 0.2 0.4 0.6 0.8 1.0
Non dimensional radial position, r/R R
Figure 4-15: Inflow ratio (at 5” below rotor) versus non-dimensional radius for the isolated 9.5” MA, 3 blade rotor
88
Chapter 5
Thrust Experiments in Hover
The ability of a ducted fan air vehicle to hover effectively is crucial to its mission.
It has been seen in Chapter 2 that the surrounding of a rotor with a duct has an effect on
the overall thrust performance in hover. It has also been seen that by varying parameters
such as rotor tip clearance and duct geometry, the thrust performance can be altered.
This chapter will continue the evaluation of ducted fans in hover through several thrust
experiments with the ducted fan wind tunnel models.
5.1 Description of Experiment
In this experiment, the thrust was measured for ducted and isolated rotor cases.
The goal was to quantify the difference in overall thrust between the two configurations.
Also several parameter variations were introduced. For the LL1 series model, the
isolated fan and ducted fan were measured and experiments included the effects of blade pitch, blade taper, and blade tip shape on the thrust production. The same 4-bladed rotors
that were used in the induced velocity experiments were used in the thrust experiment.
The effect of rotor tip clearance was studied with the 10-1 model. Three tip
clearances were investigated with the 3-bladed MA rotor and 4-bladed APC rotor as
described earlier. The tip clearances are noted as a percentage of the rotor radius. The
smallest tip clearance tested was 1%. This means that the distance between the rotor
89 blade tip and wall of the duct was 1% of the rotor radius. The other tip clearance cases
were 2% and 4%.
Finally, the effect on thrust that the opening and closing of the forward inlet vents
of model 10-2 has was determined. In this case, the thrust was measured when the vents
were closed and when the vents were open. Only the 3-bladed rotor and 1% tip clearance
were used for these experiments
5.2 Description of Facility
Two facilities were used for the thrust experiments. These included the APB and
the Hammond wind tunnel. Each facility has its own force balance device that can be
operated inside or outside of the wind tunnel. The APB force balance however is
different in style and construction than the Hammond force balance. The APB force balance is a six-component pyramid type balance and the Hammond balance is a five-
component platform type. Since each balance has its own subtleties, further discussion is
warranted on the individual aspects of the APB and Hammond force balances.
5.2.1 APB Balance
As mentioned, the APB balance has the ability to measure six components or
degrees of freedom. This balance can independently measure lift, drag, and side forces as
well as pitch, roll, and yaw moments. A force transfer schematic is shown in Figure 5-1.
As illustrated in Figure 5-1, a three-dimensional scale model can be attached to the struts
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at the balance’s focal point. The focal point is the central position where forces and
moments are being generated by the model at which the balance was designed to
measure. The balance is labeled a pyramid type because the focal point is located at the
Figure 5-1: Force transfer schematic of APB balance
tip of an imaginary pyramid formed by the structure of the balance. The struts of the balance are adjustable to a maximum of 3 feet apart. This means that the maximum
distance between the strut connection points on any model must be 3 feet. All forces and
moments are transferred through the balance struts, table, metal flexures, and eventually
load cells. The six load cells are each made up of four strain gauges in a wheatstone bridge which send the output voltage to the data acquisition system. These voltages are
then used to determine the aerodynamic loading in each direction. The process of
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converting the output voltages from the balance strain gauges to useful aerodynamic
loadings is performed through a calibration process which will be discussed next.
Other noteworthy features of the APB balance include the angle of attack, AoA,
adjustment, side slip adjustment, height adjustment system, and locking mechanism. An
angle of attack adjustment bar can be physically attached to a model as shown in Figure
5-2. This bar is attached to a lever under the balance table which can be moved up and
down via a stepper motor and controller. Using this apparatus results in overall angle of
attack adjustment of +/- 15 o.
Figure 5-2: Angle of attack adjustment bar of APB balance
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In order to change the sideslip angle of the model, a small dial located at the bottom center of the balance can be turned. As the dial is turned, gears connected to the
center shaft of the balance table also turn. This allows for the model to undergo sideslip
angles of +/- 30 o.
The entire balance rests securely on a hydraulic lift cart. The height of the balance off the floor can easily be adjusted by operating an external hydraulic pump. The balance can be raised a maximum of 3 feet off the ground. Four wheels are also attached
to the bottom of the cart. The entire mobility system allows the balance to be moved in
and out of the wind tunnel test section.
Finally, the balance features a locking mechanism. A second small dial located just below the sideslip adjustment dial can be turned to bring metal plates around the
center shaft of the balance. These plates can become tightened around the shaft and the balance table will not be able to move. Locking the balance is very important when
moving or working on the balance to ensure that the fragile flexures are not damaged.
The balance should only be unlocked during an experiment.
5.2.2 Calibrating the APB Balance
As mentioned above, a calibration is required in order to convert the output
voltages of the balance strain gauges into useful aerodynamic loading information. This
is done through the determination of an influence coefficient matrix. Consider the
following. The six forces and moments are labeled as channel numbers according to the
data acquisition system as:
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Channel 1 = Drag Force Channel 4 = Roll Moment
Channel 2 = Side Force Channel 5 = Pitch Moment
Channel 3 = Lift Force Channel 6 = Yaw Moment
The balance output can be organized into an equation as shown by equation 5-1.
Equation 5-1 states that the output voltages from the balance are equal to an influence
coefficient matrix multiplied by the forces and moments that are generated by the model.
Ch (1 V ) − Ch (1 V ) 0 Drag Ch (2 V ) − Ch (2 V ) Side 0 Ch (3 V ) − Ch (3 V )0 6x6 Lift = (5-1) Ch (4 V ) − Ch (4 V ) 0 ICM Roll Ch (5 V ) − Ch (5 V ) Pitch 0 Ch (6 V ) − Ch (6 V )0 Yaw
Note, the voltages on the left hand side of equation 5-1 are adjusted by the tare voltage
shown with a subscript zero. The influence coefficient matrix, ICM, consists of cross-
talk terms (on the off-diagonal) that are due to imperfections that exist in any balance.
For example, the drag output will be affected to some degree by each of the other loads.
Thus since the APB balance is a six-component balance, the ICM will be a 6x6 matrix of
terms. This is further illustrated by equation 5-2.
ADD ADS ADL ADR ADP ADY A A A A A A SD SS SL SR SP SY
ALD ALS ALL ALR ALP ALY ICM = (5-2) ARD ARS ARL ARR ARP ARY A A A A A A PD PS PL PR PP PY AYD AYS AYL AYR AYP AYY
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In order to obtain each term in equation 5-2, a calibration must be performed. The balance calibration can be performed inside or outside of the wind tunnel. Regardless of the location, the procedure is the same. The balance must be loaded with a known amount of weight purely in one direction. Then the voltage output from each strain gauge must be recorded. Each time successive loads are added in the pure direction, the output voltage is recorded from each strain gauge. Generally, the maximum load and load increment is determined by the estimated amount of force that is to be measured with the desired model. Finally, a plot of load versus voltage is created. The slope of this curve should be linear as each time a weight was added, the voltage should have changed accordingly. It is the slope of this curve that becomes one of the coefficients in the ICM. Thus, the ICM is simply the slopes of each calibration curve.
For example, consider a pure drag loading. A system of pulleys and bars was set up inside the wind tunnel test section at APB. A special calibration model was attached to the force balance struts. At one end of a string was the calibration model, and at the other end was a load pan. By using the pulley configuration shown in Figure 5-3, only a force in the drag direction was being applied. Note, it is important to ensure that the drag force is being applied through the focal point of the balance. A line level was used to ensure the drag force was applied parallel to the horizon. The bar that the pulley rests on must also be level in all directions. The pulley itself must also be centered with the balance focal point. Once the setup is correct, successive weights can be added to the load pan, and Figure 5-4 can be produced. Figure 5-4 shows the voltage output from each strain gauge as weights were being added in the drag direction only. The slopes were then calculated for the uploading case for all of the data series shown in Figure 5-4.
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Figure 5-3: Pulley configuration for pure drag calibration
Thus, the slope of the drag series data becomes A DD in the ICM. This is the drag channel output due to a pure drag force. The slope of the side force series data becomes A DS in the ICM. This is the side force channel output due to a pure drag force. This process is continued for the rest of data series in Figure 5-4. This will have completed all of row one in the ICM.
To complete the rest of the columns of the ICM, the same process is undertaken as was in the pure drag case. The only difference is how the pulleys are set up.
Applications of the moments usually involved applying loads at two positions. It is essential that the pulleys are setup in such a way that only pure forces and moments are applied. Pictures of the pulley setups as well as the calibration curves are shown in
Appendix B.
It is interesting to note that the dominating feature of Figure 5-4 is the drag data series. Indeed there is cross talk since the other data series are not zero, but the drag channel voltage is far greater than the others. As is shown in Appendix B.2, all the
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calibration curves feature this distinction for the pure forces and moments. Furthermore,
this will make the ICM a diagonally dominant matrix. The A DD , A SS , A LL and so on will be the highest numerical value in each of their respective rows. This is distinctive of pyramid type balances.
Drag Calibration 2
0 0 2 4 6 8 10 12
2 Drag Side Force 4 Lift Roll 6 Pitch Output Voltage [V] Yaw
8
10 Pure drag loading [lbf]
Figure 5-4: Drag channel calibration curves for APB balance
Therefore, the completed ICM is shown in equation 5-3. A history of past
calibration influence coefficient matrices is also presented in Appendix B.3.
− .0 9298 − .0 0022 .0 0089 .0 0009 .0 0013 .0 0015 − .0 0161 .1 0302 .0 0089 .0 0000 .0 0000 − .0 0051 − .0 0064 − .0 0068 − .0 4032 .0 0011 .0 0006 − .0 0001 ICM = (5-3) .0 1249 − .0 0658 .0 0139 − .0 2643 .0 0043 − .0 0063 .0 0395 − .0 0508 − .0 0025 − .0 0048 − .0 0733 − .0 0043 − .0 0606 − .0 0002 .0 0190 − .0 0005 − .0 0002 − .0 2705
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5.2.3 Hammond Balance
The Hammond balance can measure 3-components. These include lift, drag, and pitching moment. These three components however are measured through 5 load cells,
similar in design to the APB balance. Two load cells (comprised of strain gauges) are
devoted to measuring lift, and two load cells are reserved for measuring drag. The fifth
and final load cell measures pitching moment. The Hammond balance is labeled a platform-type because all of the load cells are located in the same plane. As shown in
Figure 5-5 provided by Litz[15], a scale-model can be connected to the Hammond balance between the two struts or to one strut. The struts can be adjusted up to 2 feet between them, which is the width of the Hammond wind tunnel. The forces generated by
the model are passed through the struts into the platform where all of the load cells are
located. Finally, the forces reach the load cells through small metal flexures that are
connected to the platform.
Figure 5-5: Hammond force balance layout, provided by Litz[15]
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Also shown in Figure 5-5, the Hammond balance is effectively split into two
sides, the near side and the far side. Hence, there is a near and far lift load cell as well as
a near and far drag load cell. The contribution to the total force in the near and far sides
is accounted for in the calibration and post-processing stage. Another feature of the
Hammond balance is cross-talk between each load cell channel. Due to the inherent
design of this platform balance, the load paths do not pass through a focal point of the balance. Thus pure loads will cause noticeable changes in voltage to the other load cells.
The Hammond balance also features an angle of attack adjustment. Just aft of the pitch load cell, there is a hole for the AoA adjustment rod. The rod consists of two pieces
with a pivot between them to allow for most model shapes and sizes. The rod is
connected to a stepper motor also attached to the balance. Controlled through an external
unit, the total angle of attack adjustment is +/- 20 o. An aerodynamic shielding is also placed around the adjustment rod for reduced drag in wind tunnel testing.
The balance can be used both internally and externally to the wind tunnel test
section. When inside the test section, the balance is inserted through the ceiling. The
weight of the balance is supported by a frame located above the test section. The height
of the balance can be adjusted through slots built into the frame. Set screws then hold the balance at the desired height. When the balance is located outside of the wind tunnel, it
is turned upside down from its configuration in the test section. The balance is secured by simply resting on the floor. In both locations, care should be taken to ensure that the balance is level.
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In general, there is no locking mechanism like the APB balance. The Hammond balance is always in “measuring-mode.” The load cells as well as flexures are
sufficiently protected from damage by the structure of the balance.
5.2.4 Calibrating the Hammond Balance
The introduction of the near and far side of the balance causes complexities in the
calibration. Recall that the total lift and drag is split into four load cells. The calibration procedure is nearly identical to that of the APB procedure, however the influence
coefficient matrix formulation is different. The five channels of the Hammond balance
are divided as:
Ch 1 = Near Lift
Ch 2 = Near Drag
Ch 3 = Pitch
Ch 4 = Far Lift
Ch 5 = Far Drag
In order to accommodate the near and far sides of the balance, there are two influence
coefficient matrices as shown by equations 5-4 and 5-5.
Ch (1 V ) − Ch (1 V ) 0 3x3 Lift near Ch (2 V ) − Ch (2 V ) = ICM Drag 0 near near (5-4) Ch (3 V ) − Ch (3 V )0 Pitchnear
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Ch (4 V ) − Ch (4 V ) 0 3x3 Lift far Ch (5 V ) − Ch (5 V ) = ICM Drag 0 far far (5-5) Ch (3 V ) − Ch (3 V )0 Pitch far
Notice that each ICM consists of a 3x3 matrix. Also notice that there is a near and far pitch. There is actually only one pitch load cell, therefore the near pitch and far pitch
influence coefficients are equal. The split here is for convenience. This is accounted for
in the post-processing. The near side ICM is further detailed in equation 5-6 and the far
side ICM is shown in equation 5-7.
ALL ALD ALP ICM = A A A near DL DD DP (5-6) A A A PL PD PP near
ALL ALD ALP ICM = A A A far DL DD DP (5-7) A A A PL PD PP far
Again a system of pulleys and bars can be setup inside the wind tunnel test
section or outside. In the present case, the balance was calibrated outside of the wind
tunnel. A special calibration bar was attached to the struts and angle of attack adjustment
rod. The angle of attack was set to 0 o. The pulleys were setup such that only pure forces were being applied to the balance. The voltage output from each of the five load cells was then recorded as successive loads were added to the balance.
The lift channel was calibrated by applying a pure lift force in the center of the two struts as shown in Figure 5-6. A LL , A DL , and A PL for the near ICM were found by plotting the voltage output of the near lift load cell, near drag load cell, and pitch load
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cell, respectfully, against weight loading. The slopes of these curves are inserted into the
ICM. A LL , A DL , and A PL for the far ICM were found in the same manner, only now the voltages from the far lift load cell and far drag load cell were plotted.
Figure 5-6: Pure lift calibration of Hammond balance
The drag channel was calibrated by applying a pure drag force to the center point between the two balance struts as shown in Figure 5-7. In a similar fashion to the lift calibration, the output voltage of the near lift load cell, near drag load cell, and pitch load cell were plotted against the increments in pure drag loading. The slopes of these curves became A LD , A DD , and A PD for the near ICM. The slopes of the calibration curves of the
far load cells were inserted into A LD , A DD , and A PD of the far ICM.
Finally, a pure pitching moment was applied to the balance as shown in Figure 5-
8. Again, the output voltage of the near lift load cell due to a pure pitch was plotted
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Figure 5-7: Pure drag calibration of Hammond balance
against the pitch loading increments. The slope of this curve becomes A LP for the near
ICM. The rest of the coefficients were calculated accordingly.
Figure 5-8: Pure pitch calibration of Hammond balance
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The completed influence coefficient matrices for the Hammond balance are
shown in equations 5-8 and 5-9. Notice that the A PL , A PD , and A PP coefficients are identical for both the near and far matrices. This is because there is only one pitch load cell. Also, notice there is not a diagonal dominance in the ICMs of the Hammond balance as there was with the APB balance. Caution should be taken when describing the calibration setup as an application of pure forces and moments. The Hammond balance is actually configured such that a “pure” lift will also cause a pitching moment. The same is true for the drag calibration.
Ch (1 V ) − Ch (1 V )0 .0 2706 − .0 3205 .0 0072 Liftnear Ch (2 V ) − Ch (2 V ) = .0 0267 .0 2484 .0 0009 Drag 0 near (5-8) Ch (3 V ) − Ch (3 V )0 − .0 0140 .0 6613 − .0 0139Pitchnear
Ch (4 V ) − Ch (4 V )0 .0 2691 − .0 3005 .0 0076 Lift far Ch (5 V ) − Ch (5 V ) = .0 0196 .0 3234 .0 0009 Drag 0 far (5-9) Ch (3 V ) − Ch (3 V )0 − .0 0140 .0 6613 − .0 0139Pitch far
5.3 Data Acquisition Systems
Each facility has an independent data acquisition, DAQ, system. The physical
hardware used in the APB facility is different from that used in the Hammond facility.
However, the basic setup is similar. Figure 5-9 shows the basic structure of both data
acquisition systems. At one end is the DAQ computer. This computer has all of the
software and drivers necessary to acquire and store experimental data. Inside the DAQ
computer is a type of PCI acquisition card. Since the PCI acquisition device does not
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have direct signal connectivity, a connector block must act as an interface for all of the
sensors and signals. In these experiments, the balance voltage signals as well as rotor
RPM signals are connected by BNC cables to the connector block. The rest of this
section will detail the specifics of each facility’s data acquisition system.
Figure 5-9: Basic DAQ structure
5.3.1 APB Data Acquisition System
A connection diagram of the experiment setup at APB is shown in Figure 5-10.
The software used in the DAQ computer is the National Instruments, NI, LabView
version 7.1. A program was created to control, monitor, and collect data for the entire
thrust experiment. The instructional use of this program will be explained later. The data
acquisition device used inside the computer was the NI PCI-6259 16 ch, 16 bit M-series
multi-function DAQ device. A driver was required to allow the computer to
communicate with the PCI device. Currently there are two kinds of drivers that support
NI devices. These include the Traditional DAQ and DAQmx drivers. Some newer NI
devices are only supported by the DAQ mx driver. In the case of APB, the only driver
that was installed on the computer was DAQmx, so this was the one that was used. The
105 connector block used in APB was the NI BNC 2090. The connector block was connected to the PCI device through a SH68-68 cable. The NI BNC 2090 connector block features
8 channels in the factory default setting differential, DIFF, mode. It is also capable of splitting each channel to a maximum of 16 channels in referenced or non-referenced single end mode.
Figure 5-10: DAQ setup at APB
For the thrust experiment, the connector block was set to the DIFF setting. Figure 5-11 shows the front panel switch configuration that was used. Channels 1-6 of the NI BNC
2090 are occupied by the voltage signals coming from the load cells of the force and moment balance. A strain amplifier acts as an interface between the force balance and the NI BNC 2090.
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Figure 5-11: Connector block switch configuration for hover thrust experiments
The six cables coming from the balance are connected to the strain amplifier
through an MS type connector. The entire balance can be easily disconnected from the
DAQ system just by unplugging this connector. The strain amplifier features onboard
auto and manual voltage trimming as well as gain, filter, and excitation voltage control.
During the calibration and experiment the gain was set to x10000. The low pass filter
was 1000 Hz and the excitation voltage was 5 V. It is important that these settings are
not changed during the experiment because that would alter the calibration data.
Channel 7 of the NI BNC 2090 connector block is occupied by an optical sensor signal. Normally, the optical sensor signal is also passed through a digital spectrum analyzer through a tee-connector. This allows for quick-look debugging. Depending on which model is being operated, either the Monarch optical sensor or elogger optical sensor is used. Figure 5-12 shows a connection diagram for the case of the elogger optical sensor hook up. A male BNC connector was spliced into the servo cable of the elogger optical sensor so that a voltage pulse signal could be sent to the BNC 2090. In this configuration, the motor RPM can be monitored and recorded by both the DAQ computer and an external laptop computer connected via USB to the elogger.
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Figure 5-12: Elogger hook-up for hover thrust experiments at APB
5.3.2 APB LabView Software Operation
As mentioned, the DAQ computer utilizes the LabView software for the
experiment. The balance calibration and hover thrust experiment both use a LabView program. The programs are very similar, with the thrust experiment program requiring a
few extra components. First, the hover thrust experiment LabView program will be
explained.
The filename of the code used during the experiment is
BalanceMsrwithRPM_V02.vi. The objectives of the program are three-fold. First, the
code is to acquire analog input voltage from the amplifier of the balance. This input
voltage is coming from the six load cells of the balance. Second, the code needs to
acquire an analog input signal of an optical sensor and calculate the rotating speed of the
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rotor. Finally, the code needs to store the data as an ASCII output file for ease of post- processing. The code does not require any sub vi to run but does require that the DAQmx
driver is installed in the computer. The code will run with any NI data acquisition device
that is supported by the DAQmx driver but may not with the Traditional DAQ driver.
Also, the code was written in LabView 7.1. It is possible that if the code is run on a more
recent version of LabView, it will not function properly. The code should especially not be modified and saved on a newer version of LabView. This will cause the code to not
function on version 7.1.
When the program is first opened, the front panel shown in Figure 5-13 will be
displayed. The front panel is divided into several sections. In the upper left corner is a box labeled INPUT PARAMETERS. This box also contains several parameters. The physical channel box allows for the control of what the input channels to the PCI DAQ
device are. A pull-down menu will be displayed if the arrow to the right of this box is
selected. All of the available channels for the PC will be shown. Since channels 1 - 7 are being used by the BNC 2090, they should be selected. The physical channel box should
read “Dev1/ai1:7” as shown in Figure 5-13. Dev1 denotes that device number 1 is being
used. Next the sampling rate and number of samples to read per acquisition loop can be
selected. For the hover experiment a sampling rate of 2000 Hz and 1000 samples to read
should be used. These should be entered as seen in Figure 5-13. The last input parameters allow the user to control the minimum and maximum value of the analog input voltage range. Since the voltage range of each load cell in the balance is +/- 10 V, these values should be entered as shown in Figure 5-13.
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The next section of the front panel is labeled TEST CONDITIONS. This space allows for the ambient pressure and temperature conditions to be recorded. There is also a space to make any comments about the experiment such as model configuration, angle of attack setting, or rotor type. Since everything inside the TEST CONDITIONS box gets printed to the output file, any comment that would be helpful at a later date should be recorded. Figure 5-13 also shows a sample of the TEST CONDITIONS box.
Figure 5-13: Front panel of LabView software for hover thrust experiments at APB
Moving to the middle of the screen, there are boxes labeled balance signal, laser signal, detected frequency (Hz), and no of pulse per revolution. The balance signal box is
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a graphical display of the balance voltage signals that are being acquired in real-time.
The mean voltage of each channel is also displayed at the top just to the right of the balance signal box. Under the balance signal box is the laser sensor box. The box shows
a graphical display of the acquired optical sensor voltage signal in real-time. This display
is also used to determine the quality of the optical sensor signal in addition to the
spectrum analyzer. The detected frequency box displays a numerical value of the
frequency of the optical sensor signal. This value should match with the value obtained by the spectrum analyzer. Finally, the no of pulse per revolution box corresponds to the
number of reflective surfaces applied to the rotor for the optical sensor to read. For
example, if the four bladed rotor is being used, this box should have a “4” in it. A “3”
should be used for the three bladed rotor. If the correct number is not used here, the program will not calculate the rotational speed correctly. The calculated rotational speed
in revolutions per minute, RPM, is displayed in real-time just below the average voltage
displays of each channel on the right side of the front panel.
On the right center portion of the front panel, there is the average time box. This box allows the user to control the averaging time of each measurement. For the hover
thrust experiments, an averaging time of 10 seconds was used. This means that during 10
seconds, all the signals are acquired, averaged, and stored. Just below the average time box are the inputs of voltage and current. These values are entered during the
experiment. For the 10 series models, the voltage current values are read from the
elogger software through the laptop computer and not the power supply’s digital display.
These values will change for each rotor RPM case. These values are also recorded to the
output file. These input boxes are one of the only differences in the calibration code as
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will be seen later. The Start Average and STOP buttons are used during the experiment.
When the program is running and data is ready to be acquired, the “Start Average” button
should be pressed. While the program is acquiring data, a status bar will move across the
screen just under the “Start Average” button. It is important that while the data are being
acquired, the balance is not touched. Vibrations can be transmitted to the balance load
cells which will cause an error in the data. The “Start Average” button will need to be pressed each time a data point or test case is desired to be recorded. When the
experiment is complete, the “STOP” button should be pressed. This will cause the program to stop running. At the bottom of the front panel is a box labeled STORED
DATA. This box displays the mean values of the data for all the test cases that were
acquired. This is the data table that gets stored in an ASCII output file.
Finally, in the upper left corner of the panel, just below the edit menu bar, there is
a white arrow. This arrow should be selected when the user is ready to run the program.
Essentially, this button should be pressed after the user has selected the correct input parameters and made all the necessary comments. Once the arrow is selected another
window will pop up that asks the user to create a filename for the output file. It is helpful
to label the filename with a .txt extension. The location of the output file can also be
selected at this time. Once the experiment is over, the file can be accessed at the location
defined in this step.
The filename of the code used for calibration is BalanceMsr for
Calibration_V02.vi. This code is identical in layout and operation to the previous code
used for thrust measurement with the exception of a few parts. Figure 5-14 shows the
front panel for the calibration code. Because, there is no use for an optical sensor during
112 calibration, the laser sensor display window, detected frequency, no of pulse per revolution, and rotational speed boxes are all removed. The main difference however is in the parameter input boxes on the left center portion of the panel.
Figure 5-14: Front panel of LabView software for the calibration of the APB balance
Recall before, there were only two parameter input boxes labeled voltage and current. For the calibration code, these are changed to correspond to the six channels of the force balance. The amount of load being applied during in the calibration is simply input to these boxes. For example during the drag channel calibration, the amount of load applied in the pure drag direction is recorded in the Drag parameter box. After the
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“Start Average” button is pressed and the data is acquired, the next load can be applied
and recorded in the parameter box space. The “Start Average” button is pressed again
and the process is repeated until all the loadings have been recorded for each channel.
Typically, a separate output file is created for each channel.
5.3.3 Hammond Data Acquisition System
The data acquisition system used with the Hammond force balance is very similar
to that used at APB. There are some slight differences in the LabView software of the
DAQ computer to accommodate five channels of the balance instead of six. Also, the
amplifier used between the balance and connector block is different from APB. The
main difference however, is in the DAQ hardware that is used in the Hammond facility.
The Hammond facility uses legacy National Instruments hardware. The data
acquisition device used inside the computer was the NI PCI-6036 16 ch, 16 bit E-series
multi-function DAQ device. The driver used was the Traditional DAQ. This means that
a different LabView program had to be created for the experiments in Hammond since
the code at APB utilizes the DAQmx driver. Also, the NI BNC 2120 connector block
was used. This block features 7 channels of input in the ground source reference.
Channels 1 - 5 were occupied by the load cells of the force balance. Again, an amplifier
was used as an interface between the force balance and NI BNC 2120.
Figure 5-15 shows the amplifier used in the Hammond facility. The amplifier
features digital displays which have menu controls for gain settings. There is also a tare button on the display. Notice that in Figure 5-15, there are six displays, although the
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sixth display is not connected to anything. Five cables from the force balance are
connected to the front of the amplifier through standard MS type connectors. Then BNC
cables connect each channel of the amplifier to the NI BNC 2120. The gain settings are
not as easily modified as in APB. The gains were set so that 1 volt (as seen on the
amplifier digital display) was equal to 1 unit of load. For example, if 2 pounds were
applied in lift, the far lift channel would display 1 volt and the near lift channel would
display 1 volt. Together, the two lift channels add to 2 volts which corresponds to 2 pounds.
Figure 5-15: Type Caption Here
Channel 6 of the NI BNC 2120 was occupied with the optical sensor signal.
Again, a digital spectrum analyzer was used to quick-look the signal before entering the
DAQ computer.
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5.3.4 Hammond LabView Software Operation
Again, the LabView software is used in the Hammond facility DAQ computer. A program was written for taking thrust measurements as well as calibrating the balance.
Essentially, the programs written for the Hammond facility are identical to that of the
APB facility. In the case of the thrust measurement, the Hammond program still acquires balance voltage signals, optical sensor voltage signals, and creates output files in ASCII format. The key difference is that the Hammond program is setup for a 5 channel force and moment balance instead of 6. Also, the Hammond programs are written for a
Traditional DAQ driver. The Hammond DAQ computer runs LabView 7.1.
The name of the program used for taking thrust measurements in the Hammond facility is BalanceMsrwithRPM_V01.vi. The front panel of this program is shown in
Figure 5-16. As shown in Figure 5-16, the box labeled channels under the INPUT
PARAMETERS section requires a different selection. Since only six channels are used for the hover thrust experiment in the Hammond facility (5 for the balance and 1 for the optical sensor) the device box should read “1” and the channels box should read “1:6.”
This program also allows the user to define the buffer size. In these experiments, the buffer size was chosen to be 10000. The Hammond program also allows the user to select the voltage range of each channel of the DAQ. For the balance channels, a range of 0 - 10 V was selected and the optical sensor was set at +/- 10 V. The only other key difference in this program is in the mean voltage display of the balance channels. As shown in Figure 5-16, the TEST CONDITIONS/comments box remains along with the
116 balance signal display, laser sensor display, and stored data table. The operation of this program is also identical to that of the APB program detailed earlier.
Figure 5-16: Front panel of LabView software for hover thrust experiments at Hammond
The filename of the code used for calibrating the Hammond balance is
BalanceCal_497K_V02.vi. Again, this code runs identical to its APB counterpart. The
two differ in appearance by the input parameters section on the center right side of the
front panel. Seen in Figure 5-17, there are only two input parameters for the Hammond balance calibration. The amount of the pure load in the lift, drag, or pitch direction can be entered there.
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Figure 5-17: Front panel of LabView software for the calibration of the Hammond balance
5.4 Post-Processing
Post-processing in general refers to the steps that must be taken in order to convert the output voltages that were acquired into useful information in terms of forces and moments. The post-processing step also includes the calculation of non-dimensional aerodynamic coefficients and the visual display of these quantities. Post-processing is crucial in order to make any interpretations on the aerodynamic phenomena associated
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with these experiments. It has already been seen that the balance calibration is utilized in
this step. The next two sections discuss the specifics of the post-processing steps
required for the APB and Hammond facilities.
5.4.1 APB Post-Processing
Referring back to equation 5-10, the LHS are the voltages (corrected for tare)
from the force balance and the ICM on the RHS is determined from the calibration.
Ch (1 V ) − Ch (1 V ) 0 Drag Ch (2 V ) − Ch (2 V ) Side 0 Ch (3 V ) − Ch (3 V )0 6x6 Lift = (5-10) Ch (4 V ) − Ch (4 V ) 0 ICM Roll Ch (5 V ) − Ch (5 V ) Pitch 0 Ch (6 V ) − Ch (6 V )0 Yaw
Now the only unknown in equation 5-10 is the force and moment column on the RHS.
These are the forces and moments that correspond to the output voltages acquired during the experiment. In order to solve for the forces and moments, the LHS of equation 5-10 must be multiplied by the inverse of the influence coefficient matrix. Thus equation 5-11 shows how the output voltages of the balance are converted into dimensional forces and moments.
−1 Drag Ch (1 V ) − Ch (1 V ) 0 Side Ch (2 V ) − Ch (2 V ) 0 Lift 6x6 Ch (3 V ) − Ch (3 V )0 = (5-11) Roll ICM Ch (4 V ) − Ch (4 V )0 Pitch Ch (5 V ) − Ch (5 V ) 0 Yaw Ch (6 V ) − Ch (6 V )0
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5.4.2 Hammond Post-Processing
Recall that the Hammond balance has two channels for drag, two for lift, and one for pitch. Equations 5-12 and 5-13 show the relationship between the balance output voltage and the calibration matrices.
Ch (1 V ) − Ch (1 V ) 0 3x3 Lift near Ch (2 V ) − Ch (2 V ) = ICM Drag 0 near near (5-12) Ch (3 V ) − Ch (3 V )0 Pitchnear
Ch (4 V ) − Ch (4 V ) 0 3x3 Lift far Ch (5 V ) − Ch (5 V ) = ICM Drag 0 far far (5-13) Ch (3 V ) − Ch (3 V )0 Pitch far
In order to solve for the near forces and far forces, again the LHS of equations 5-12 and
5-13 are multiplied by the respective inverse influence coefficient matrices. Thus equations 5-14 and 5-15 show how the near and far forces are converted from balance voltages to dimensional forces.
−1 Lift near 3x3 Ch (1 V ) − Ch (1 V ) 0 Drag = ICM Ch (2 V ) − Ch (2 V ) (5-14) near near 0 Pitchnear Ch (3 V ) − Ch (3 V )0
−1 Lift far 3x3 Ch (4 V ) − Ch (4 V )0 Drag = ICM Ch (5 V ) − Ch (5 V ) (5-15) far far 0 Pitch far Ch (3 V ) − Ch (3 V )0
Unique to the Hammond balance, the near and far forces and moments are components that make up the total quantities. In order to compute the total force and moment, the components are combined as in equations 5-16, 5-17, and 5-18.
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Lifttotal = Liftnear + Lift far (5-16)
Dragtotal = Dragnear + Drag far (5-17)
Pitch + Pitch Pitch = near far (5-18) total 2
The total lift force is found by simply adding the calculated near and far lift. It is
assumed that the lift is equally distributed between the near and far load cells. The same
is true for the total drag. Since there is only one pitch load cell on the balance, the
calculated near pitch and far pitch are averaged to find the total pitch. Essentially, the
near pitch is equal to the far pitch.
5.4.3 Non-Dimensionalization of Forces and Moments
Most of the results from the hover thrust experiment were non-dimensionalized for comparison purposes. The thrust coefficient was introduced as shown in equation 5-
19.
L CT = 2 (5-19) 5.0 ρVTip S DA
It should be noted that in equation 5-19, the lift force, L, is replaced with the horizontal force, in the case where the model in mounted in the 90 o position. The density in
equation 5-19 is the local ambient density in the facility where the experiment was
conducted. The rotor tip speed was calculated as shown in equation 5-20.
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RPM V = 2π R R (5-20) Tip 60 D
RPM R in equation 5-20 is simply the rotor RPM. Also R D is the radius of the duct. This
dimension was used since several rotors of varying tip clearance were used during the
experiment. The radius of the inner duct is a constant parameter. S DA in equation 5-21 is
defined as a ducted fan disk area.
2 S DA = πRD (5-21)
Finally, the rotor tip speed was non-dimensionalized by equation 5-22.
V M = Tip (5-22) Tip a
The speed of sound, a, was calculated as shown in equation 5-23.
a = γRT (5-23)
The ratio of specific heat constants, γ , of 1.4 was used in this calculation. Also, R is the
universal gas constant and T is the ambient temperature of the experimental facility.
5.5 Experiment Setup
In all cases, the protective screens were placed around the balance when exposed.
The hover thrust experiment was performed inside the APB wind tunnel test section for model 10-2 with the vents open and closed. In the case of the Ford Fan and LL1 experiment, the Monarch optical sensor was used to measure the rotor RPM. In the 10 series experiments, the elogger optical sensor was used to measure rotor RPM and the elogger micro temperature sensor was used to monitor motor temperature. The next
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section discusses more detail of the individual setup of each model during the hover
thrust experiment.
5.5.1 LL1-1 Setup
The hover thrust experiment with model LL1-1 was performed in both the
Hammond and APB facilities. In both cases, the same test conditions were observed.
Both balances were also located outside of the wind tunnel for the experiment. The
Hammond balance was rested on the floor next to the Hammond wind tunnel. It should be noted that the balance was inverted from its position inside the wind tunnel test section, where the struts hang from the ceiling. The test conditions included mounting the isolated rotor upside down and at the 90 o position. The ducted rotor was also
positioned in these two configurations. Both the tapered tip (11” pitch) and square tip
(12” pitch), 14.5” 4-blade rotors were used in the APB facility. It was expected that the
rotor with the higher pitch distribution would produce a greater amount of force. The
difference in thrust production due to blade taper and blade tip shape was also examined.
Only the tapered tip, 4-blade rotor was used in the Hammond facility. This was done as a
means to create a second set of data as well as compare the two facilities. Figures 5-18
and 5-19 show the various configurations in the Hammond and APB facilities.
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Figure 5-18: Model LL1 setup for hover thrust experiments on Hammond balance
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Figure 5-19: Model LL1 setup for hover thrust experiments on APB balance
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5.5.2 10-1 Setup
The focus of the 10-1 hover thrust experiment was to measure the difference in thrust produced when rotors of varying tip clearance were used. Blade number and blade pitch distribution were also varied. The ducted 10-1 model was mounted upside down on the APB balance. The angle of attack adjustment bar was utilized to position the plane of the rotor at 0 o to the horizon. The angle was easily checked using a digital level across the flat surface of the duct exit. Three tip clearance cases were used with the 4-blade
11”x6” APC rotor. These included 1%, 2%, and 4% of the rotor radius. Two tip clearance cases, of 2% R and 4% R, were used for the lesser pitched 3-blade 10”x5” MA rotor. The tip clearances were adjusted by shaving material off the tips of each blade until the desired radius was obtained. The tip clearances were checked and documented using feeler gauges.
5.5.3 10-2 Setup
The hover thrust experiment of model 10-2 was performed with the APB balance inside the test section of the APB wind tunnel. The model was again mounted upside down and leveled to 0 o using the angle of attack adjustment bar. Since the test section is
5 feet wide by 3.25 feet tall, it was expected that the walls could have an affect on the measurement. The model should be out of ground effect however since it was mounted such that the distance between the test section ceiling and duct exit was greater than 2 rotor diameters. The goal of this experiment was to measure the difference in thrust produced when the forward vents of model 10-2 were closed and when the vents were
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open. The vents were closed both on the outside and inside of the duct using a special
adhesive tape designed for sealing leaks in ventilation ducts. This tape was ideal because
it would remain fairly stiff when exposed to airflow. The opening and closing of the
forward vents in illustrated in Figure 5-20. Only the 3-blade 10”x5” MA rotor with a tip
clearance of 1% R was used for this test.
Figure 5-20: The forward vents of model 10-2 can be opened (Left) and closed (Right)
5.6 Experiment Procedure
Regardless of which balance was used, a voltage tare must be performed before any data are acquired. Essentially, this involves adjusting the output voltage of each load cell of the balance as close as possible to 0 V. With the APB strain amplifier, this is simply done by pressing the auto-tare switch for each channel. The voltage can further be adjusted by turning the manual trim knobs located just below the auto switches. With the Hammond balance, the auto-tare button on each display unit of the amplifier can be pressed. In both cases, the LabView software displays a real-time wave form and numeric of the voltage of each load cell which can assist in the tare process.
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The same Xantrex 40-18 DC power supply was used as before. Recall that the power supply was directly connected to the motor the LL1 model. For the LL1 model,
the input voltage was increased until the current limit of the power supply was reached.
The current was allowed to change based on the amount of loading placed on the motor by the rotor. First, the tapered tip, 4-blade rotor was used. In the Hammond facility experiments the input voltage was ranged from 4 to 7, and back to 4 V in increments of
0.5 V. A maximum of 7.3 V was reached. The unloading case was performed to check the hysteresis of the measurement as well as provide a second set of data. Overall, the rotor RPM ranged from 1700 to 2600. When the LL1 model was tested at the APB facility a new set of longer wires was used to connect the motor to the power supply.
This effectively increased the resistance of the input power circuit and thus changed the input voltage range. At APB the LL1 input voltage ranged from 4 to 9.5 V in increments of 0.5 V. This resulted in a rotor RPM range of 1500 to 2800 RPM. Although, the rotor
RPM was not substantially increased, the new wires allowed for more data points to be acquired. The voltage was also stepped down from 9.5 to 4 V for the same reasons as before. With the square tip (higher pitched) 4-blade rotor, a maximum input voltage of 9
V was used. This resulted in a rotor RPM range of 1400 to 2500. The hover test matrix for the LL1 model is shown in Table 5-1.
The servo tester was introduced with the 10 series models as a better means of
rotor RPM control. For model 10-1, the input voltage was set at 10 V for the duration of
the experiment. As adjustments were made with the servo tester, the input current was
allowed to change with the loadings being applied to the motor by the rotor. Resistance
changes in the power wires due to temperature fluctuation caused the input voltage to
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Table 5-1: Test Matrix for Hover Thrust Experiments with model LL1
Facility Rotor Duct ON/OFF Position Power Supply Voltage [V]
90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5 ON 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5 14.5"x11" 4 blade tapered tip 90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5 OFF 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5
APB
90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9 ON 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9 15.5"x12" 4 blade square tip 90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9 OFF 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9
90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.3 ON 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.3 14.5"x11" Hammond 4 blade tapered tip 90 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.3 OFF 180 o 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.3
Notes: 1. 90 o position is thrust measured by the drag load cell 2. 180 o position is thrust measred by the lift load cell but with the model exhausting toward the ceiling 3. All input voltages were loaded to maximum value and then back down in increments as shown vary within 5% of the original 10 V setting. The first rotor used was the 11”x 6” 4-blade
APC rotor. The servo tester was used to vary the rotor RPM from 2500 to 5000 in increments of 500 RPM. Two measurements at 0 RPM were also made to check the tare of the balance. However, instead of conducting a hysteresis test, the exact test case of
2500 to 5000 RPM was repeated for a total of 5 times or trials. This allowed for the repeatability of the measurement to be understood. An average of the data from all 5 trials was taken. Five trials of data were acquired for each of the tip clearances used with
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the 4-blade APC rotor. Identical procedures were followed for the 10”x 5” 3-blade MA
rotor. However, since the overall pitch is lower for the MA rotor, the maximum rotor
RPM was increased to 6000. Table 5-2 shows the hover test matrix for model 10-1.
Table 5-2: Test Matrix for Hover Thrust Experiments with model 10-1 Rotor Duct ON/OFF Tip Clearance Rotor RPM [%R] [RPM]
1 2500, 3000, 3500, 4000, 4500, 5000 11"x 6" 4 blade square tip ON 2 2500, 3000, 3500, 4000, 4500, 5000
4 2500, 3000, 3500, 4000, 4500, 5000
2 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 10"x 5" ON 3 blade square tip 4 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000
Notes: 1. Input power supply voltage was set at 10 V 2. 5 trials of each tip clearance case were performed for repeatability and averaging
A similar test procedure was performed with the 10-2 model. Since the same 3- blade MA rotor was used, an identical range of rotor RPM was obtained. However, instead of taking an average from 5 sets of data, only 3 repeat trials were performed. The hover thrust test matrix for model 10-2 is shown in Table 5-3.
Table 5-3: Test Matrix for Hover Thrust Experiments with model 10-2 Rotor Vents OPEN/CLOSED Rotor RPM [RPM]
OPEN 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 10"x 5" 3 blade square tip CLOSED 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000
Notes: 1. Input power supply voltage was set at 10 V 2. 3 trials of each vented case were performed for repeatability and averaging 3. A single tip clearance of 1% R was used 4. Test was performed inside APB wind tunnel test section
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5.7 Hover Thrust Experiment Results
Figure 5-21 shows the hover thrust experiment results for the LL1 model. The results were non-dimensionalized and presented as thrust coefficient versus rotor tip
Mach number. The thrust produced by the isolated 4-blade rotor with a tapered tip was measured by the lift load cell as well as the horizontal load cell on the APB balance. In both cases, the thrust increases with rotor RPM. For the isolated rotor case, a maximum thrust coefficient of just over 0.035 was obtained at a tip Mach number of 0.16. The thrust measured through the horizontal load cell is higher than that measured through the lift load cell.
14.5" Rotor,Tapered Tip *APB Balance 0.050
0.045 LL1 duct 0.040 T 0.035
0.030 Isolated rotor 0.025
0.020
0.015 without Duct, Thrust = Lift Thrust coefficient, C without Duct, Thrust = Horizontal 0.010 with LL1 Duct, Thrust = Lift 0.005 with LL1 Duct, Thrust = Horizontal 0.000 0.000 0.050 0.100 0.150 0.200
Rotor tip Mach number, M Tip
Figure 5-21: Hover thrust coefficient versus rotor tip Mach number for 14.5” tapered tip, 4 blade rotor with and without LL1 duct
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The addition of the duct, also shown in Figure 5-21, appears to have brought the measurements made by each load cell closer together. However, in the ducted rotor case, the thrust measured by the lift load cell was higher than that measured by the horizontal load cell. It should be mentioned that the lift load cell of the APB balance has a maximum load capacity of 100 + pounds. Since model LL1 produced a dimensional thrust on the order of 2 pounds, the lift load cell of the APB balance measured data on the low end compared to its maximum capacity. This could be improved however, by applying a lower range of loads for the pure lift calibration.
Despite this inconsistency between the load cell measurements, Figure 5-21 suggests an expected result between the ducted and un-ducted cases. At a maximum tip
Mach number of 0.16, a thrust coefficient of about 0.045 was produced. This is almost a
30% gain in thrust coefficient from the isolated rotor case.
Figure 5-22 shows the hover thrust results for the 14.5”x 12” 4-blade rotor with square tips. Again the isolated rotor thrust coefficient measured through the horizontal load cell is higher than that measured through the lift load cell for all tip Mach numbers.
A maximum thrust coefficient of about 0.045 is obtained at a tip Mach number of 0.15.
However, notice that the discrepancy between the two load cell measurements is less than with the tapered tip rotor. These observations are likely due to the fact that the square- tipped rotor has a higher loading distribution than the tapered tip rotor. The addition of the LL1 duct around the 14.5”x 12”, square tip, rotor is also shown in Figure 5-22. The difference in measurement between each load cell for the ducted rotor case is less than
2%. For the ducted case, a maximum thrust coefficient of 0.06 is produced at a tip Mach
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number of 0.15. For the square-tipped rotor with a higher pitch distribution, this
represents over 30% thrust augmentation with the addition of the duct.
14.5" Rotor,Square Tip *APB Balance 0.070
0.060
LL1 duct T 0.050
0.040 Isolated rotor
0.030
without Duct, Thrust = Lift
Thrust coefficient, C 0.020 without Duct, Thrust = Horizontal with LL1 Duct, Thrust = Lift 0.010 with LL1 Duct, Thrust = Horizontal
0.000 0.000 0.050 0.100 0.150 0.200
Rotor tip Mach number, M Tip
Figure 5-22: Hover thrust coefficient versus rotor tip Mac h number for 14.5” square tip, 4 blade rotor with and without LL1 duct
Finally, in order to compare two different facilities, the hover thrust experiment
with the tapered tip, 4-blade rotor was repeated with the Hammond balance. Figure 5-23
shows the thrust coefficient versus rotor tip Mach number for the ducted rotor case only
measured from both the APB and Hammond balances. In both facilities, the thrust
measured by the lift load cells is greater than that measured by the horizontal load cells
over all tip Mach numbers. However, it appears that the entire data set obtained from the
Hammond balance is about 10% less than the APB balance data. This deviation seems to
decrease as the rotor tip speed increases. It should also be noted that the two balances are
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completely different from each other in design and function. Also, the data acquisition
systems are different between the facilities. Given these differences, the hover thrust performance of the LL1 ducted rotor model is repeatable and comparable.
14.5" Rotor,Tapered Tip 0.05
0.045
0.04
T 0.035
0.03
0.025
0.02
0.015 with LL1 Duct, Thrust = Lift, Hammond Thrust coefficient, C with LL1 Duct, Thrust = Horizontal, Hammond 0.01 with LL1 Duct, Thrust = Lift, APB 0.005 with LL1 Duct, Thrust = Horizontal, APB
0 0 0.05 0.1 0.15 0.2
Rotor tip Mach number, M Tip
Figure 5-23: Hover thrust coefficient versus rotor tip Mach number for 14.5” tapered tip, 4 blade rotor with LL1 duct measured at Hammond and APB
Now that the thrust performance has been compared for an isolated and ducted
rotor, the effect of varying the tip clearance inside the duct was then studied. Martin[5]
showed that as the tip clearance became larger, the thrust of the ducted fan decreased.
Similar experiments were performed with the 10-1 model in hover at the APB facility.
Initially, three tip clearance test cases with the 4 blade APC rotor were used. The results
of this experiment are shown in Figure 5-24. The thrust coefficient was plotted versus
rotor tip Mach number for each tip clearance case. As seen in Figure 5-24, the thrust
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coefficient increases with tip Mach number for each case as expected. Also, the smallest
tip clearance case of 1% yields the highest amount of thrust over the entire tip Mach
number range when compared to the other tip clearances. The thrust produced by the 2%
tip clearance rotor is greater than that of the 4% tip clearance rotor for the majority of tip
Mach numbers. However, at tip Mach numbers 0.13 and below, the measured thrust produced with the 4% tip clearance rotor is higher. This result was not expected and
generally is not thought to be of physical significance.
APC 9.5" Square Tip Rotor, 4 blade 0.044 *Model 10 1 T 0.04
0.036
1% R TC 0.032 Thrust coefficient, C 2% R TC 4% R TC 0.028 0.05 0.1 0.15 0.2 Rotor tip Mach number, M Tip
Figure 5-24: Tip clearance effect with model 10-1, 9.5” APC, 4 blade rotor
To further investigate the tip clearance effect, a repeat experiment was performed
with the 3 blade, MA rotor. Although the MA rotor is of lesser pitch than the APC rotor
(and thus a lower thrust force will be measured), the reduction to 3 blades may have a
different effect. Figure 5-25 shows the results of the tip clearance study with the 3 blade
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MA rotor at just the 2% and 4% tip clearance cases. The expected result is shown such
that the 2% rotor produced a larger amount of thrust over all tip Mach numbers when
compared to the 4% rotor. Thus the anomaly shown with the 4 blade rotor of Figure 5-25
appears to be due to experimental error. Overall, the tip clearance studies support the
claim that higher thrust can be produced as the rotor tip clearance becomes smaller.
MA 9.5" Square Tip Rotor, 3 blade 0.03 *Model 10 1
0.025 T
0.02
0.015
0.01 Thrust coefficient, C 0.005 2% R TC 4% R TC 0 0 0.05 0.1 0.15 0.2 0.25
Rotor tip Mach number, M Tip
Figure 5-25: Tip clearance effect with model 10-1, 9.5” MA, 3 blade rotor (Non- dimensional)
For future reference in Chapter 6, the dimensional thrust as a function of rotor RPM is presented for model 10-1 with 2% and 4% tip clearances as shown in Figure 5-26.
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MA 9.5" Square Tip Rotor, 3 blade 1.2 *Model 10 1
1
0.8
0.6 Thrust [lbs] 0.4
0.2 2% R TC 4% R TC 0 0 2000 4000 6000 8000 Rotor RPM
Figure 5-26: Tip clearance effect with model 10-1, 9.5” MA, 3 blade rotor (dimensional)
The final hover thrust experiment included the determination of the “vent effect”
on model 10-2. This was accomplished by measuring the changes in thrust that occur
when the forward vents are opened and closed in hover. Figure 5-27 shows the results of
this experiment. Interestingly, when the vents are opened, the thrust coefficient increases
fairly linear with increases in rotor tip Mach number. However, the opening of the vents
causes a nearly constant 15% decrease in the thrust coefficient compared to when the
vents were closed. The open vents effectively “bleed” air away from the duct exit.
Although the slipstream does not become distorted enough to cause random thrust
fluctuations, the opening of the vents does reduce the amount of air that is ultimately
expelled at the duct exit. Thus the purpose of such vents on the forward face of the duct
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is unclear in terms of hover thrust performance. The vent concept with model 10-2 will be further explored in a series of forward flight experiments inside the wind tunnel at
APB in Chapter 6.
MA 9.5" Square Tip, 3 blade rotor, 1%R TC 0.035 *Model 10 2
0.03 T 0.025
0.02
0.015
0.01
Thrust coefficient, C Vents Closed Vents Open 0.005
0 0 0.05 0.1 0.15 0.2 0.25
Rotor tip Mach number, M Tip
Figure 5-27: Comparison of thrust coefficient for model 10-2 with MA rotor when forward vents are open and closed
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Chapter 6
Forward Flight Experiments (Wind Tunnel)
The final experimental effort with these ducted fan models was conducted inside a wind tunnel. Recall that the primary forward flight regime of interest was that of edgewise flight and shallow angles of attack. This is the likely environment that ducted lift fans would be exposed to. In this chapter, crucial elements to the success of such a vehicle in edgewise flight will be examined including lift, drag, and pitching moment performance. The idea of momentum drag as explained in Chapter 2 will be revisited.
Also the effect of opening the forward vents on forward flight performance will be explored.
6.1 Description of Experiment
In the forward flight experiment, all 6 forces and moments were measured for the ducted and isolated rotor cases. These included drag, side force, lift, roll, pitch, and yaw.
The primary goals were to examine the difference in aerodynamic performance between the isolated and ducted rotor, compare the performance of differently shaped ducted fans, and to investigate the “vent” effect over a wide range of wind speeds. While all this was being done, a vast experimental database was being formulated.
Only the 10 series models were tested in the wind tunnel. For all cases, only the 3 bladed master airscrew 9.5”x 5” rotor was used with a tip clearance of 1% R. The same
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rotor RPM range of 2500 to 6000 was also used. Finally, all the models were operated
over a shallow range of angle of attack equal to -3o, 0 o, and +3 o. Recall that 0 o is defined
as edgewise flow. Baseline measurements were taken with the isolated rotor model. The
model was operated over a range of four wind tunnel speeds including 0, 10, 25, and 66
ft/sec.
Model 10-1 was initially installed in the test section without the rotor. The duct
only, was subjected to wind speeds ranging from 0 to 70 ft/sec. This was intended to
gain a better understanding of the forces and moments produced by the duct itself. When
the 3 bladed rotor was installed, a rotor RPM sweep was performed in wind speeds
including 0, 10, 25, 50, 62, and 66 ft/sec.
The duct of model 10-2 was similarly tested in the wind tunnel with the rotor
removed. Changes in performance due to different duct shape could be obtained. With
the rotor installed, forces and moments were also measured over a range of wind tunnel
speeds with the forward vents opened and closed.
6.2 Description of Facility
The entirety of the forward flight experiments were performed in a single wind tunnel facility. As with any wind tunnel, a series of calibrations were required to perform the experiment. Forces and moments were measured during the experiment again using the 6 channel APB balance. Changes needed to be made to the configuration of the balance and a calibration check was performed. All of these issues are discussed next.
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6.2.1 APB Wind Tunnel
The major wind tunnel facility at the The Pennsylvania State University is located in the academic projects building, APB. Swan[16] and Germanowski[16] provide details of the facility. The APB wind tunnel is a closed-loop type and has overall dimensions of
75 feet long by 25 feet wide. Figure 6-1 shows a top-view diagram of the wind tunnel.
The wind tunnel is mostly comprised of wood except for the long return diffuser section which is made of steel.
The wind tunnel is powered by a single 300 hp, 480 volt Elektrim electric motor
(model sgm 449-68) which is located external to the second corner of the tunnel. The motor itself is capable of operating at a maximum rotational speed of 1200 RPM. An
ABB ACS 800 adjustable frequency AC inverter drive is used to control the amount of power being supplied to the motor and thus wind tunnel speed. An eight bladed fan of roughly 6 feet in diameter is connected to the motor through a direct drive shaft. The use of the AC inverter drive can effectively vary the speed of the fan from 0 to 1200 RPM.
The drive shaft passes through an opening of a cascade of airfoils in the second corner.
The airfoils have a chord of 8 inches with 2 inch spacing between each successive airfoil.
The cascade serves to smoothly guide the re-circulating airflow around the corner and into the fan. Just downstream of the fan is a set of 13 stator vanes that act to reduce the swirl components exiting the fan. A center body, shown in Figure 6-2, is also affixed to the stator vanes to facilitate in smoothing the exit flow of the fan.
The return diffuser of exit-to-inlet area ratio 2.2:1 transitions the tunnel cross section from circular at the fan to rectangular at the entrance of the third corner. The
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Figure 6-1: Top view schematic diagram of APB wind tunnel
142 airflow is again guided through the third and fourth corners by a set of airfoils or vanes.
The vanes in the third and fourth corners however consist of two sets each with 6 inch chords and 3 inch spacing. Brophy[18] notes that the second set of vanes in the third corner are stiffened to prevent vibrations caused by vortex shedding. The cross section of the tunnel between the third and fourth corners remains constant and rectangular.
Figure 6-2: APB wind tunnel fan and stators
Just after the fourth corner is a series of screens. The first screen consists of fine grid spacing. This screen is used for preventing small particles such as dirt and dust from recirculating around the tunnel. Immediately downstream of the dust screen, the tunnel cross section expands in a 2:1 area ratio. The expanded region is called the settling chamber. Inside the settling chamber, there is another screen made of a perforated plate.
This screen acts to reduce large scale air flow separation. The third set of screens located in the settling chamber is a square grid honeycomb. The square grid size is 0.375 x 0.375 inches. The honeycomb grid extends 6 inches in the downstream direction and helps to
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reduce swirl components. The final set of screens vary in grid mesh size and are used to
destroy any shear layers that exist in the airflow. All of the screens located in this section
of the tunnel are used to reduce the overall turbulence intensity and improve the
steadiness of the airflow inside the test section.
Continuing downstream of the settling chamber, the tunnel cross section contracts
in a 9.3:1 area ratio. This contraction section is the last section upstream of the test
section. The contraction section accelerates the air flow coming from the settling
chamber. The corners of the contraction section feature tapered fillets in order to
maintain low turbulence and reduce vorticity. Finally, two static pressure ports are
located at separate locations inside this section. As will be described later, the pressure ports are used for calibrating the wind tunnel and determining what the test section
velocity is.
The wind tunnel test section features a rectangular cross section. The test section is 5 ft wide, 3.25 ft tall, and extends 6.1 ft downstream. The corners of the test section are also filleted to reduce corner vorticity. The downstream exit of the test section diffuses at an angle of 1.2 o to reduce the static pressure gradient and boundary layer
growth throughout the section. Also at the exit is a breather port to help maintain
atmospheric pressure inside the test section. The walls of the test section are composed
of removable plexiglass windows. This allows for easy access to the model during setup.
Eight sets of lighting, both white and black, line the inside of the test section. A custom
removable floor was fabricated to accommodate the APB force and moment balance.
Holes were cut, symmetric about the tunnel center line, into the floor to accommodate
models ranging in width of approximately 8 inches to 30 inches. A single slot of length
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17 inches and width 1.5 inches was also cut in the floor to allow for sufficient travel of
the angle of attack adjustment bar. The test section ceiling features a slot for a traversing pitot static probe which is used during tunnel calibration. A thermocouple sensor is also positioned through a hole in the ceiling which measures changes in test section
temperature and thus density during an experiment.
The final section to complete the closed loop wind tunnel is labeled the first
diffuser. This section acts to decrease the dynamic pressure by slowing the air flow as it
re-enters the fan region. From beginning to end, the first diffuser section expands with an
area ratio of 1:1.9. The cross section also transitions from rectangular to octagonal
downstream of the first corner. Similar to the turning vanes of the second corner, a set of
8 inch chord airfoils are located in the first corner. A vane spacing of 2 inches is also
used. A wire mesh net is located just upstream of the first corner in order to collect any
debris that may separate from the model during an experiment. Three maintenance
hatches are located in this section. Inspections of the wire mesh net, corner vanes, and
wind tunnel fan can be made at these hatches.
Located in the center of the wind tunnel structure, adjacent to the test section, is the wind tunnel control station. The speed control for the wind tunnel is located there.
The data acquisition equipment is also located on the raised wooden platform. A section of the platform was modified in order to position the APB force and moment balance under the center of the test section. Since the platform is at the same height as the test section, the experimenter can have easy access to the model.
In 1994, Brophy[18] performed a series of experiments in the APB wind tunnel facility to measure the tunnel flow quality. A maximum test section velocity of 220 ft/sec
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was achieved. As shown in Figure 6-3, the turbulence intensity was found to increase
with wind tunnel speed. The same year Medina[19] performed experiments with an S805
airfoil. These measurements taken in the APB facility were compared to measurements
taken in a low-speed wind tunnel at Delft University using the same airfoil. The Delft
wind tunnel has an excellent reputation so the very good agreement between the results
demonstrated the high quality flow in the PSU wind tunnel.
Figure 6-3: Measurement of turbulence intensity of APB wind tunnel, Brophy[18]
6.2.2 Calibrating the Test Section
The wind tunnel test section speed can be determined with a pitot-static probe inserted in the center of the test section. The pitot-static probe measures the difference between the stagnation pressure and static pressure, known as the dynamic pressure.
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Using Bernoulli’s equation, the test section velocity can be calculated as a function of
dynamic pressure as shown in equation 6-1.
2 P U = (6-1) ρ
The P in equation 6-1 is the test section dynamic pressure. The density, ρ , is the density inside the test section and can be calculated by the equation of state shown in equation 6-2.
P ρ = (6-2) RT
The static pressure, P, and temperature, T, in equation 6-2 are also the conditions inside the test section. For the current experiment however, the ambient pressure recorded from a barometer external to the test section is used in place of the static pressure. It is assumed that the diffusion angle of the test section and the breather port allow for the ambient pressure to be used. Finally, the R in equation 6-2 is the ideal gas constant.
Thus, the pitot-static probe can measure the test section velocity. However, the existence of the model in the vicinity of the probe can distort the measurement. Recall that there are two static pressure taps located in the contraction section of the wind tunnel. A comparison can be made between the pressure difference in the contraction section and the dynamic pressure inside the empty test section. Then with the model installed during the experiment, the pitot-static probe can be removed and the test section velocity can be obtained by the pressure difference measured in the contraction section.
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This process generally describes the test section calibration and is further illustrated by
the diagram shown in Figure 6-4 of the APB wind tunnel. The contraction section, test
Figure 6-4: Diagram of APB wind tunnel for calibration
section, and diffuser of the APB wind tunnel are illustrated. As shown, there is a static pressure tap at station 1 which has a tunnel cross-sectional area of A 1 and there is a
second static pressure tap at station 2 of cross-section area A 2. Both stations 1 and 2,
which are upstream of the test section, also have a velocity of V 1 and V 2 respectively.
Pcont is the difference in pressure between stations 1 and 2. Finally, U is the test section
velocity. From Bernoulli’s equation, the pressure at each station in the contraction
section can be written as shown in equation 6-3.
1 1 P + ρV 2 = P + ρV 2 (6-3) 1 2 1 2 2 2
Equation 6-3 can be rearranged to find the difference in pressure between stations 1 and
2. This is shown in equation 6-4.
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1 P − P = P = ρ[V 2 −V 2 ] (6-4) 1 2 cont 2 2 1
Also, from the conservation of mass, the volumetric flow rate can be compared at each
station as shown in equation 6-5.
A1V1 = A2V2 (6-5)
When equation 6-5 is inserted into equation 6-4, a relationship between the pressure
difference in the contraction section and the cross-sectional areas of each section can be
obtained. This is shown in equation 6-6.
2 1 A P = ρV 2 1− 2 (6-6) cont 2 2 A 1
Finally, if it is assumed that the velocity at station 2 of the contraction section, V 2, is
equal to the test section velocity, U, a relationship between the test section dynamic pressure and pressure difference in the contraction section can be obtained as shown in
equation 6-7.
1 P P = ρU 2 = Cont 2 A 2 (6-7) 1− 2 A 1
Therefore, if the cross-sectional areas are known at the locations of each pressure
tap in the contraction section, the dynamic pressure in the test section can be found and
hence the test section velocity can be calculated. However, equation 6-7 does not
normally hold due to boundary layer growth inside the wind tunnel or incorrect area
estimations. To correct for this, a linear relationship is made between the dynamic
149 pressure inside the test section and the pressure difference inside the contraction section
as shown by equation 6-8.
P = K PCont (6-8)
Notice that equation 6-8 is the equation of a straight line. A pitot-static probe measures
the dynamic pressure at the center of the empty test section over a range of wind tunnel
speeds. The pressure difference in the contraction section is also recorded at each speed
condition. A plot is then made of dynamic pressure versus contraction pressure
difference. The plot should be linear and the slope simply becomes K in equation 6-8.
The test section was calibrated once in the summer of 2008 and once in the spring
of 2009. The K value in 2008 was found to be 2.764 and K value in 2009 was found to be 2.753. The velocity inside the wind tunnel test section is then calculated as shown in
equation 6-9.
2K P U = cont (6-9) ρ
Three Validyne pressure transducers (model DP15-24) were used during the wind
tunnel operation. One transducer was used to measure the static pressure difference in
the contraction section. A second and third pressure transducer was used to measure the
total and static pressure in the test section separately. Each of these transducers also
needed to be calibrated to convert a voltage output into units of pressure. A hand held pump was used to apply a load to the transducers. An oil manometer was also used to
measure the amount of load being applied to the transducers. A plot was then made of
the known applied pressure load versus output voltage from the transducer. After
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applying a linear regression to the curve the slope was recorded as the transducer
sensitivity. The transducer sensitivities for the contraction section pressure, total pressure, and static pressure in the test section were found to be 0.01526 psi/volt,
0.05101 psi/volt, and 0.05161 psi/volt respectively. A separate transducer was used to
measure the static pressure inside the test section so that changes with velocity could be
monitored. A quadratic relationship between test section static pressure and pressure
difference in the contraction section was found as shown in equation 6-10.
2 Pstatic = .2 0431 Pcont + .0 6178 P − .0 0002 (6-10)
A calibration was also performed with the K-type thermocouple which measures
temperature changes inside the test section. The thermocouple was calibrated using cold
and boiling water. Also a mercury thermometer was used to measure the temperature of
the water. The voltage output from the thermocouple was recorded with changes in the
water temperature. Again, a linear curve fit was applied to the plot of measured
temperature versus thermocouple voltage output. The thermocouple sensitivity was
measured to be 18.55 oF/volt.
6.2.3 APB Balance
The APB balance was used again to measure forces and moments during the
forward flight wind tunnel tests. The balance was wheeled under the test section such
that the side force load cell and accompanying structure fit into the pre-cut slot in the
acquisition platform. The balance was aligned with the center line of the test section with
two plumb bobs. The balance was shifted fore and aft to ensure that the focal point of the 151 balance was in the center of the wind tunnel test section. Initially the balance was raised
with the hydraulic cart so that the model was located in the vertical center of the test
section; however, this configuration caused greater hysteresis error in the balance voltage
signals. Thus it was decided to only raise the balance 8 inches off the floor and support
the balance with two steel beams at each corner as shown in Figure 6-5. Since the position of the model was now closer to the floor of the test section, it was decided to perform the experiments with the model mounted upside down.
Figure 6-5: APB balance supported under the test section
In order to save time, the APB balance was not calibrated again. Instead, a simulated known load was applied to the balance. By applying a known load to the balance, the previous calibration and hence influence coefficient matrix could be
“checked.” Known loads were applied in the lift, drag, and pitching moment directions simultaneously as shown in Figure 6-6. Generally, the magnitudes of the simulated loads
152 are determined by the estimated loads that will be applied during the experiment. The hover thrust experiments showed a maximum measurement of around 1.5 lbs through the lift channel. However since the supply of known weights was limited, increments of 2 lbs were applied to the balance in the lift direction up to a maximum of 10 lbs.
Increments of 1 lb up to 5 lbs were applied to the drag channel. It should be noted that negative drag and lift forces were being applied. The pitching moment arm for the simulated load was 8 inches or 0.67 feet. Two lb increments were added up to 10 lbs so that a maximum positive pitching moment of 7 lb-ft was applied. In all three simulated load cases, the amplifiers were zeroed with 1 lb weights added to each load pan. The same LabView program used for the balance calibration was used for the simulated load.
Figure 6-6: Simulated load setup
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Figure 6-7 shows the lift channel results of the simulated lift load test. The data
series is plotted as applied lift force versus run number. The run number is simply a
count on the number of acquisitions obtained. Notice that the square data points
represent the actual amount of weight being added to the lift channel and the “x” data points is the amount of lift as measured by the balance. The measured lift appears to be
roughly 10% less than the actual applied load. More troublesome however is the slope of
the measured lift curve between 0 and 2 lbs. At around 1 lb of measured lift, the slope of
the curve abruptly changes. This suggests that at the lower end of applied weight to the
lift channel, the balance measurement could be erroneous and inconsistent. This is a problem because it is expected that the models will produce amounts of lift in this region.
0.00 0 2 4 6 8 10 12
2.00
4.00
6.00 Lift [lbf]
8.00
10.00 Sim Lift w/o weights on table
Actual Lift Load 12.00 Run No.
Figure 6-7: Initial APB lift channel simulated load results
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It was suggested that this balance phenomena was due to a shifting or loosening of flexures during the transfer of the balance under the wind tunnel test section and not because of an old calibration. It had also been suggested that applying a pre-loaded weight to the lift channel could force the lift flexures to “reseat” themselves. Thus, two methods were used correct this problem. First, the lift channel flexures were all tightened in place. Then several weights were added to the center of the balance table as shown in
Figure 6-8.
Figure 6-8: Preloaded weights added to the center of the APB balance
After the amplifiers were re-zeroed with the preloaded weights, simulated lift loads were again applied. It was determined that the optimal amount of preloaded weight was 7 lbs. The new simulated lift curve is shown in Figure 6-9. Again, the lift force is plotted against acquisition run number. Now however, there are three data series. Two of the series represent the measured lift with and without the preloaded weight. The third series is the actual amount of lift added to the balance. As seen in Figure 6-9, not only did the addition of the preloaded weights to the center of the balance table reduce the
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overall error, but it also removed the phenomena between 0 and 2 lbs. The error was
actually seen to be smallest in this low lift region. The error grows to 7% at the
maximum applied load. Overall, the measured lift was still less than the actual lift being
applied.
2.00
0.00 0 2 4 6 8 10 12 2.00
4.00
Lift [lbf] 6.00
8.00
Sim Lift w/o weights on table 10.00 Sim Lift w weights on table Actual Lift Load 12.00 Run No. Figure 6-9: APB lift channel simulated load after added weights to center of table
The simulated negative drag and positive pitching moment are shown in Figures
6-10 and 6-11. The dimensional force or moment is again plotted versus run number.
The measured load is also plotted with the actual applied load. Both the simulated drag
and pitching moment appear to be in agreement with the actual applied loads. Error of
around 1% exists between the two. It was concluded from the results of the simulated
load test that the balance did not have to be recalibrated. The previous influence
coefficient matrix was deemed suitable and the forward flight experiments could proceed.
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0.00 0 2 4 6 8 10 12
1.00
2.00
3.00 Drag [lbf]
4.00
5.00 Simulated Drag Load Actual Drag Load 6.00 Run No.
Figure 6-10: Negative drag simulated load on APB balance
8.00 Simulated Pitch 7.00 Actual Pitch Load
6.00
5.00
4.00
Pitch [lb ft] 3.00
2.00
1.00
0.00 0 2 4 6 8 10 12 Run No.
Figure 6-11: Positive pitching moment simulated load on APB balance
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6.3 Changes to the Data Acquisition System
With the additional sensors required for a wind tunnel test, some changes needed
to be made to the data acquisition system previously used for the hover thrust
experiments. Particularly, the connecter block needed to modified to allow for 9 input
channels. A new amplifier was also introduced for the wind tunnel sensors. Finally, the
hover thrust LabView program was modified to include visual displays of wind tunnel
speed as well as extra quick look capability. Additionally, all of the data acquisition
equipment was now located on the wooden platform control station next to the wind
tunnel test section.
6.3.1 Hardware Configuration
The core of the data acquisition system remains unchanged. The same National
Instruments PCI 6259 16 channel, 16 bit M series multi-function DAQ device is located
in the data acquisition computer. The BNC 2090 connecter block is also attached to the
PCI device through the SH68-68 cable. The key change in the system is how the BNC
2090 connecter block is configured. Previously in the hover thrust experiment, only 7
channels were used. Six were needed for the force and moment balance and one was
needed for the rotor RPM. For the wind tunnel experiments, an additional two channels
are required for the pressure transducer signal in the contraction section and the
thermocouple signal in the test section. In total, there are now 9 channels which must be
used. It is impossible to use 9 channels in differential mode on a 16 channel connecter block. Therefore, part of the connecter block had to be split into differential mode and
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referenced single end mode. It was decided to maintain differential mode on channels 1
through 7. These would be used for the balance and rotor RPM sensor. The channel 0
and channel 8 inputs were separated by selecting the reference single end mode. Channel
0 was used for the pressure difference in the contraction section and channel 8 was used by the thermocouple in the test section. Figure 6-12 shows a diagram of the switch
configuration on the back of the BNC 2090 connecter block.
Figure 6-12: BNC 2090 configuration for wind tunnel test at APB
The same strain amplifier was used for the force and moment balance. Shielded
cable coming from each of the six load cells of the balance connected to the amplifier via
a MS type connecter. BNC cables then connected the amplifier to the BNC 2090. The
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same 10,000x gain and 5 V excitation voltage was used as in the hover thrust experiment.
The low pass filter was also set to 1000 Hz.
Since only the 10 series models were tested in the wind tunnel, the elogger RPM
sensor was used to monitor the rotor RPM. As mentioned earlier, a male BNC connecter
was spliced into the servo cable of the RPM sensor. A wire was soldered from the pin of
the BNC connecter to the white signal wire of the servo cable. A second wire was also
soldered from the ground post of the BNC connecter to the black wire of the servo cable.
The rotor RPM signal was again monitored through BNC cable by a digital spectrum
analyzer and the LabView software. An external laptop computer was also used via USB
connection with the elogger device. A micro-temperature sensor was also placed at the bottom of the motor and connected to the elogger. The motor temperature was also
monitored with the laptop computer. Figure 6-13 shows a screenshot of the elogger
control panel.
Figure 6-13: Elogger control panel
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A second acquisition system (computer, PCI device, connecter block) was permanently located on the wooden platform control station. This system was in addition
to the one being used with the APB force and moment balance. The second DAQ system
contains an amplifier for which the pressure difference in the contraction section and
temperature in the test section signals are connected to as shown in Figure 6-14. There is
also a digital display on the amplifier that can be used to view the instantaneous voltages
of either signal. Switches located under the display select which channel is displayed.
The pressure difference signal was connected to channel 5 of the amplifier and the
temperature signal is connected to channel 6 of the amplifier. There is a manual tare
knob for the pressure difference signal as well as a gain selection knob. If the span (gain)
is changed during the experiment, the tunnel will have to be recalibrated so caution
should be taken.
Figure 6-14: APB wind tunnel amplifier
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The wind tunnel signals were then connected to a second BNC 2090 connecter block as shown in Figure 6-15. Channel 5 of the tunnel amplifier (pressure signal) was
connected to channel 6 of the tunnel connecter block. Channel 6 of the tunnel amplifier
(temperature signal) was connected to channel 7 of the tunnel connecter block. It is at the
tunnel connecter block that the signals are transferred to the balance connecter block
through T-joints and BNC cables. Thus the pressure signal from channel 6 of the tunnel
connecter block gets transferred to channel 0 of the balance connecter block. The
temperature signal from channel 7 of the tunnel connecter block is transferred to channel
8 of the balance connecter block. Only the balance connecter block needs to be
configured for referenced single end mode. A connection diagram of the entire balance
data acquisition system is shown in Figure 6-16.
Figure 6-15: APB wind tunnel connector block
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Figure 6-16: APB DAQ connection diagram for wind tunnel experiments
6.3.2 Wind Tunnel LabView Software
The LabView software for the wind tunnel experiments is very similar to the program used for the hover thrust experiments. In fact, the wind tunnel code is a
modified version of the hover thrust code. The file name of the wind tunnel LabView
code is Windtunnel Measurement_V02.vi. The program objectives remain the same with
one addition. In addition to acquiring and storing the voltage signals from the APB balance and RPM sensor, the program can now calculate the wind speed at the test
section of the wind tunnel by acquiring voltage signals from the tunnel. The new wind 163 tunnel program again requires that the DAQmx driver be installed on the computer and that LabView version 7 is used. A view of the front panel is shown in Figure 6-17.
Figure 6-17: APB wind tunnel LabView program front panel
Notice that the layout has a similar flow to the previous hover thrust program.
The key changes to the layout include updated boxes in the INPUT PARAMETERS section, pressure difference & temperature display tab, and air speed calculation. In the
INPUT PARAMETERS section, there are now separate selections for the differential channels and single ended channels. Since the six channels of the balance and the rotor
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RPM signal are differential, the input of “Dev1/ai1:7” should be selected under the
differential physical channel box as shown in Figure 6-17. The voltage range remains the
same as +/- 10 V. Just below are the boxes pertaining to the single ended channels.
Recall that the pressure difference in the contraction section is single ended and located at
channel 0 on the balance connector block. The temperature signal from the test section is
also single ended and is channel 8 on the balance connector block. Thus “Dev1/ai0,
Dev1/ai8” must be typed into the single end physical channel box as shown in Figure 6-
17. The voltage range for these channels is +/- 5 V. Also as before, the sample rate is
2000 Hz and the number of samples to read is 1000 for the wind tunnel experiments.
The TEST CONDITIONS section remains unchanged, however care should be
taken when entering the ambient pressure. This user defined value of pressure is utilized by the LabView program in calculating air density inside the test section. The air density
in turn is used for calculating the test section velocity. Note that the units of the ambient pressure are mbar. If incorrect units are used, the wind speed calculation will be
incorrect. The ambient temperature and comments are simply recorded in the output file.
The ambient temperature is not used in the wind speed calculation.
Another new feature to the wind tunnel program is the pressure difference &
temperature display tab shown in Figure 6-18. During the experiment, this window
shows a graphical display of the voltage signals for the pressured difference in the
contraction section and the temperature inside the test section. This can be used as a
quick-look debugging window. Note under the window there are two boxes. The first
one on the left is denoted as “voltage of CH0.” This box displays a numerical value of
the instantaneous mean voltage of the pressure difference signal. Immediately to the
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right of this, there is a “zero voltage for pressure transducer for PD” box. This box serves
as a second option for taring the pressure difference signal. Ideally, the pressure signal
should be zeroed with use of the manual zero knob on the tunnel amplifier. However,
sometimes the voltage drifts away from zero. If this occurs, the drifted value of voltage
can be entered into the “zero voltage for pressure transducer for PD” box as described
above. The LabView program will then automatically subtract the entered zero voltage
value from the measured pressure difference voltage.
Figure 6-18: Pressure difference & temperature display of APB LabView program
The last major change in the program is the air speed calculation window, located at the upper right of the front panel. This window is shown in Figure 6-17. The first item displayed in the window is the actual measured pressure difference in the contraction section of the wind tunnel converted from units of volts into units of pressure including pounds per square inch, psi, and pascals, pa. This calculation is made by LabView using the slope information obtained through tunnel calibration. The tunnel calibration data is manually entered into the LabView program outside of the front panel and can be modified for future experiments. This step will be discussed shortly. The next box in the
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airspeed window is a numeric display of the test section temperature in oF. Again,
LabView uses the thermocouple calibration data to convert the voltage output into temperature. The test section temperature shown there is also used in the calculation of the test section air density which is displayed directly below in units of kg/m 3. Recall that the test section air density is used in the calculation of test section airspeed. Finally, the calculated test section airspeed is numerically displayed in both units of m/s and ft/sec. All of the values displayed in this section are the instantaneous values. They all change with respect to each other.
As mentioned, the tunnel calibration information must be manually input into the
LabView program in order for the airspeed to be correctly calculated. Over time, the
calibration may change, so it is necessary to be able to change those parameters in the program. As seen in Figure 6-19, the numerical values are simply typed into the block
diagram screen of the LabView code. If the calibration values need to be modified, they
can just be re-entered in the block diagram. The program will have to be saved again
however to run. The program must be saved in LabView version 7.1 in order to work properly and it is suggested that a different filename be used.
The rest of the front panel remains very similar to the hover thrust program. The
quick-look display for the balance as well as rotor RPM signals still exists. Also, the
instantaneous average balance voltage and calculated rotor RPM appear on the right side
of the panel. An averaging time of 10 seconds was again used for the wind tunnel
experiments. Since the 10 series models all use the elogger device, the voltage and
current displayed through the elogger software is now entered into the voltage and
current input boxes of the LabView program. Previously, the voltage and current of the
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Figure 6-19 : Block diagram of the APB wind tunnel LabView program: Pressure transducer calibration (Left), test section calibration constant K (Right)
DC power supply was entered there. It is assumed that the input power displayed through the elogger is more accurate than that displayed by the power supply since the length of wire between the model and elogger is shorter. The stored data table is also updated with the extra input channels such as pressure difference and thermocouple voltage. The output file also contains the calculated airspeeds. Finally, the wind tunnel program is operated in the same manner as the hover thrust program. Initially the run button (white arrow) is pressed and a filename for the output file is recorded. When the program is running, the “start average” button is used to acquire data and the “stop” button is used to cease running of the program.
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6.4 Post-Processing
An identical post-processing procedure was completed for the wind tunnel experiments as was done for the hover thrust experiments. It was still necessary to convert the balance output voltages into dimensional forces and moments using the inverse influence coefficient matrix obtained through the balance calibration. In addition to non-dimensionalizing with rotor tip speed, the wind tunnel data was also non- dimensionalized with wind speed. Also a new external wetted ducted fan area was calculated. In order to subtract the drag of the balance struts and model support structure a special tare procedure was completed. The “true” drag coefficient of the models could then be obtained. The entire post-processing phase was aided by the creation of an excel code. The next few sub-sections discuss the details of the steps involved in the post- processing of the wind tunnel experiment data.
6.4.1 Non-dimensionalization of Forces and Moments
The six-component balance was fully utilized for the forward flight experiments.
Three forces were measured during the experiment including drag, side force, and lift.
Three moments were also measured including roll, pitch, and yaw. Each of these
components was then used to calculate aerodynamic coefficients. Each of the forces was
non-dimensionalized with wind tunnel speed and external wetted ducted fan area with the
exception of lift coefficient. The measured lift was non-dimensionalized with both wind
tunnel speed and rotor tip speed for comparison. Equation 6-11 shows the definition of
the drag coefficient before any tare was applied.
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Dm CD = 2 (6-11) 5.0 ρVWT S DF
The Dm in equation 6-11 is the measured drag of the model including the balance struts
and support structure. These extra components will be removed non-dimensionally later.
The density, ρ , is again the local ambient density. VWT is the wind speed inside the test
section and SDF is the external wetted area as defined by equation 6-12.
S DF = πDcD (6-12)
The external wetted area is found by multiplying the circumference of the duct, πD , by
the duct chord, cD. Recall that the duct chord is defined as the straight line distance between the leading edge lip and duct exit.
Both the side force and lift are non-dimensionalized in a similar manner as shown by equations 6-13 and 6-14.
S ()CS WT = 2 (6-13) 5.0 ρVWT S DF
L ()CL WT = 2 (6-14) 5.0 ρVWT S DF
Notice the WT subscript on equation 6-14. This is to denote that the lift force is non-
dimensionalized with wind tunnel speed. As mentioned earlier, the lift force was also
non-dimensionalized with rotor tip speed as shown by equation 6-15. Two definitions of
lift coefficient were used to interpret the results which will be shown later.
L ()CL Tip = 2 (6-15) 5.0 ρVtip S DF
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The moment coefficients were defined in a very similar manner. The extra length
term appearing in the denominator is the radius of the duct, R D. The pitching moment was also non-dimensionalized with rotor tip speed for the same reason as the lift force. eqautions 6-16 through 6-19 show the definitions of the moment coefficients.
l ()Cl WT = 2 (6-16) 5.0 ρVWT S DF RD m ()Cm WT = 2 (6-17) 5.0 ρVWT S DF RD
m ()Cm Tip = 2 (6-18) 5.0 ρVtip S DF RD n ()Cn WT = 2 (6-19) 5.0 ρVWT S DF cD
Finally, the wind tunnel speed was non-dimensionalized with rotor tip speed. The
advance ratio, , was defined as shown by equation 6-20.
V = WT (6-20) Vtip
6.4.2 Tare Drag Procedure
During the experiment, the balance struts and model supports are exposed to the
airflow inside the test section. Although they are designed to produce little drag, these
structures external to the model do contribute to the overall drag force being measured by
the balance. Typically, a procedure is performed so that the drag produced by everything
except the model itself can be removed from the final measurement. If done correctly,
the final total drag is only that which is being produced by the model. With the 10 series
171 models, such a tare drag procedure was performed inside the APB wind tunnel. The same could be done for the other five components of the balance, but it is assumed that the support structure will only substantially contribute to the drag measurement.
The model was removed from the test section and only the support structure remained as shown in Figure 6-20. Then the drag was measured with the balance over a range of wind tunnel velocities. The range of wind tunnel velocities was chosen to include the same velocities that were used during the forward flight experiment with the models. In this case, the range of velocities included 0 to 70 ft/sec.
Figure 6-20: Tare drag setup for ducted fan models in APB wind tunnel
The same LabView program used during the forward flight experiment was used to acquire the balance signal at each test condition. Once the drag was measured over the desired speed range, the procedure was repeated and an average was taken. For these experiments, the tare drag procedure was repeated three times and an average between the trials was taken. A plot of drag force, in units of pounds, versus wind tunnel speed 172
was then created as shown in Figure 6-21. A 2 nd order polynomial was then fit over the data curve. The equation of the 2 nd order polynomial gives the tare drag force as a
function of wind tunnel speed as shown in equation 6-21.
−5 2 Dragtare = 8x10 VWT + .0 0003VWT − .0 0004 (6-21)
struts and rods only
0.45 0.40 0.35 y = 8E-05x 2 + 0.0003x - 0.0004 0.30 R2 = 1 0.25 0.20 Drag [lbf] 0.15 Trial 1 Trial 2 0.10 Trial 3 0.05 average 0.00 0.00 20.00 40.00 60.00 80.00 Wind Tunnel Speed [ft/sec]
Figure 6-21: Tare drag versus wind tunnel speed for the ducted fan models in the APB wind tunnel
Thus the tare drag force can be calculated for any wind tunnel speed. Finally, the tare
drag coefficient was defined by equation 6-22 for the ducted fan models.
Dtare ()CD tare = 2 (6-22) 5.0 ρVWT S S
The area SS is defined as the total wetted area of the model support structure including the balance struts. One final step is applied to the tare drag coefficient before it is subtracted
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from the overall drag measurement. This step is completed in the post-processing code
and will be explained next.
Since the isolated rotor model uses a different support structure, a second tare procedure was performed. As shown in Figure 6-22, the entire isolated rotor structure
including the rods and hub were mounted in the test section. The only part that was
removed was the rotor itself. Hence, after the tare procedure is performed, only the drag
of the rotor will remain.
Figure 6-22: Tare drag setup for isolated rotor model in APB wind tunnel
The same procedure as described above was used to measure the isolated rotor
tare drag. The plot of drag force versus wind tunnel speed for the isolated rotor model is
shown in Figure 6-23. The tare drag of the isolated rotor model as a function of wind
tunnel speed is displayed as equation 6-23.
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2 (Dragtare )isolated = .0 0002VWT − .0 0001VWT + .0 0009 (6-23)
hub, struts, and rods only
1.00 0.90 0.80 y = 0.0002x 2 - 0.0001x + 0.0009 0.70 R2 = 1 0.60 0.50
Drag [lbf] 0.40
0.30 Trial 1 0.20 Trial 2 Trial 3 0.10 average 0.00 0.00 20.00 40.00 60.00 80.00 Wind Tunnel Speed [ft/sec]
Figure 6-23: Tare drag versus wind tunnel speed for the isolated rotor setup in the APB wind tunnel
6.4.3 Post-Processing Code
A post-processing code was created with multiple objectives in mind. Using the
LabView output file, the post-processing code needed to be able to quickly convert the balance voltage signals into dimensional forces and moments. Simple mathematical
operations should then be performed to the data to produce non-dimensional aerodynamic
terms. It was also desired to have the code plot the aerodynamic data as a function of both rotor RPM and advance ratio for quick-look purposes. Any inconsistency in the
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experimental data could immediately be identified and corrective actions could be taken
to ensure the quality of the data. Additionally, the code needed to be able to average
multiple trial runs of the experiment as done in previous experiments. The final output of
the code would be a table of the experiment parameters including all dimensional and
non-dimensional data with corrections for tare drag. That table could then be used to
create final plots for interpretation. Lastly, it was desirable for the post-processing code
to be easily modified for alternate configurations or be able to work with different force balances altogether. For example, the code should be able to be modified to work with
the Hammond balance output. All of these requirements made the code ideally suited for
the Microsoft Excel environment. Thus the wind tunnel post-processing code was
created in Excel 2003. The code does not require any macros to function properly and
will run on either windows or mac versions. The current code works for any six
component force and moment balance. The code is divided into three sections. There is
a quick-look graphing section, raw data input section, and averaging section.
Figure 6-24 and Figure 6-25 outline the post-processing code in block diagram
format. One nice feature of using Excel for such a purpose is that once the proper
equations are defined in their respective locations, the LabView output file data can
simply be inserted into a blank space in the spreadsheet. The post-processing code can
then quickly calculate everything it needs to. Therefore, the initial step in the post- processing code is to obtain the LabView output file.
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Text File output from Labview - experiment comments “zero” Ch. 1 Voltage V1,i - V 1,0 - run number voltages - motor voltage A Ch. 2 Voltage V2,i - V 2,0 - motor current The initial condition voltage - ch. 1-6 voltage of balance must be subtracted from Excel Ch. 3 Voltage V3,i - V 3,0 - ch. 0, ch. 8 of DAQ successive voltages of each - rotor RPM channel Ch. 4 Voltage V4,i - V 4,0 - venturi pressure difference [psi] * i = 0:n - TS temperature [deg F] Ch. 5 Voltage V5,i - V 5,0 - TS density [kg/m 3] n = number of rotor rpm changes. - TS velocity [m/s] Ch. 6 Voltage V6,i - V 6,0 - Ch. 0-Ch.8 STD - rpm, PD, temp, density, U STD Convert voltages to forces Notes: and moments APB balance (IICM) x (“zero” voltages) Ch. 1 = Drag Ch. 2 = Side Force The inverse influence Ch. 3 = Lift coefficient matrix must be Ch. 4 = Rolling Moment multiplied by the array of Ch. 5 = Pitching Moment zeroed voltages. Ch. 6 = Yawing Moment D V ,1 i −V 0,1 S V −V This code is applicable for any model AoA ,2 i 0,2 and any wind tunnel speed. L 6 x 6 IICM V ,3 i −V 0,3 = B R V ,4 i −V 0,4 The final output is a table of dimensional & P V −V non-dimensional quantities averaged between ,5 i 0,5 any number of trials for a set AoA and wind Y V ,6 i −V 0,6 tunnel speed.
Figure 6-24: Block diagram of APB wind tunnel post-processing code
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Figure 6-25: Continued block diagram of wind tunnel post-processing code
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An example of the output file is shown in Figure 6-26. As seen, everything displayed in the output table of the LabView program is saved to the output file. This includes all the experiment comments, run number, motor voltage, motor current, all the voltage signals from the DAQ as well as there standard deviations, rotor RPM, pressure difference in the contraction section, test section temperature, test section density, and finally test section velocity.
Figure 6-26: Example LabView output file from APB wind tunnel test
The LabView output file data must be manually inserted into the raw data box of the post-processing code as seen in Figure 6-27. The comments can be pasted in separately next to the raw data. The next step in the code is to remove the initial
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Figure 6-27: Raw data input section of APB wind tunnel post-processing code
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unloaded test case voltage from each balance channel. This step zeros the voltage for the
0 rotor RPM test case. It also subtracts the zero load case from each successive test case.
As seen in the lower right corner of Figure 6-27, this is done for all 6 channels of the
force and moment balance.
Next, the code multiplies the inverse of the influence coefficient matrix by the
matrix of zeroed voltages. The inverse influence coefficient matrix, IICM, can be found
at the bottom of the averaging section of the post-processing code as shown in Figure 6-
28. If a new balance calibration is performed, the new IICM can just replace the old
matrix. As long as the new matrix overlaps the exact position of the old one, the
calculations will be automatically updated. The calculated dimensional and non-
dimensional forces and moments are first tabulated just below the raw data input in
Figure 6-27 in the box labeled “measured forces and moments & non-dimensional
quantities.” The underlying equations for this section actually consist of the matrix
multiplication between the IICM and balance voltages. Caution should be used however,
since an absolute value is applied to the side force, lift, pitch, and yaw calculations. This
was done so that the magnitudes of the values would be positive. This could especially
effect the interpretation of the directions of the forces and moments. Also in this section
are displays of the rotor RPM, rotor tip speed, wind tunnel speed, and advance ratio.
Continuing with the block diagram, the next step includes an if statement. If multiple trial runs were completed during the experiment, then the process of manually inserting the LabView output file data is repeated. This is done for as many trials that were performed. Currently, the post-processing code is setup for 3 trials. Notice in
Figure 6-27, there is a Trial 1 signifier. There is a separate area signified by each trial
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Figure 6-28: Averaging section of the APB wind tunnel post-processing code
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number located below the trial 1 data set. The equations remain identical for each trial
and the same IICM is used. The only difference between each trial is the raw data that is
inserted. Once all of the trial data has been manually inserted, dimensional and non-
dimensional plots are created in the quick-look graph section as shown in Figure 6-29.
There are a total of 12 plots in this section. The data from each trial are plotted
superimposed over eachother.
Finally, the output table shown at the top of Figure 6-28 is compiled in the
averaging section of the code. The calculated dimensional and non-dimensional
quantities are averaged between each trial. For example, the measured drag coefficient before tare is applied is averaged between each trial according to equation 6-24.
nt
∑()CD j j=1 (6-24) ()C = D avg nt where j = trial number
and nt = total number of trials
Note that some of the entries in the averaged table, contain a “#DIV/0!”. This simply
informs the user that those certain entries contain a 0 in the denominator. This occurs
when the quantity is non-dimensionalized with a 0 wind speed or 0 rotor tip speed. A
correction is made to the averaged drag coefficient to account for the tare drag as
described earlier. The average tare drag coefficient is first tabulated for each average
wind speed. The corrected drag coefficient is then found using equation 6-25.
D S C = C − tare * S ()()D corrected D avg 2 (6-25) 5.0 VWT S S S DF
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Figure 6-29: Quick-look graphing section of APB wind tunnel post-processing code
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Notice that the tare drag coefficient is multiplied by the area ratio S S/S DF . This is so that the two drag coefficients can be subtracted properly since the measured average drag coefficient is non-dimensionalized with S DF and not S S.
Thus the converted balance output, averaged over multiple trials and corrected for tare drag, is obtained. Even though each step was described in detail above, it is noted again that all this is completed directly after the raw data is manually inserted from the
LabView output file. Also note from Figure 6-27 that there are multiple tabs at the bottom of the screen corresponding to different model configurations. Typically, the
entire test matrix for a single wind tunnel speed is contained in one excel spreadsheet.
Each spreadsheet can then be used as a clone for the next wind speed. The file naming
structure for the post-processed data is shown by equation 6-26. The name of the model
(6-26)
being tested is identified in the filename followed by the name of the rotor being used.
The number of blades is also listed there. The wind tunnel speed is also identified.
Finally the version number and author’s initials of the file are listed at the end. For
example the file name D10-2_MA10-3_U25vb_LMM.xls contains the post-processed
data for the 10-2 ducted fan model using the master airscrew rotor. The rotor has a 10”
diameter and 3 blades. The wind tunnel speed tested was 25 ft/sec. The current version
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of that file is “b,” and the author is LMM. In addition, this file will contain all the angles
of attack that the model was operated at for a wind speed of 25 ft/sec. In this particular
case the “vents open” and “vents closed” data will also be contained in the file.
6.5 Experiment Setup
In all cases, the 10 series models were mounted upside down to the force and moment balance inside the test section of the APB wind tunnel. Any un-necessary gaps or holes in the test section were sealed with the stiff HVAC adhesive tape. Since all the
10 series models use the same type of brushless electric motor, the elogger setup was the same for each model. A detailed model setup is shown in Figure 6-30. Power is supplied to the model by a DC power supply capable of variable voltage and current. The maximum voltage obtainable is 40 V and the maximum current is 18 A. As shown in
Figure 6-30, two wires of gauge 12, one red and one black, are connected from the power supply to the elogger. A second set of two wires then act as an intermediate connection with the electronic speed controller, ESC. Finally, the three wires coming from the ESC are connected directly to the electric motor of the model. Recall, that the servo tester is connected to the ESC and allows for manual control of the rotor RPM. Also, two sensors are connected to the elogger. The motor temperature sensor is plugged into the “Tmp 1” port of the elogger. The sensor itself is fastened to a non-rotating part of the electric motor of the model. The optical RPM sensor is plugged into the “RPM” port of the elogger. It is crucial that the optical RPM sensor is connected to this port only.
Otherwise, the sensor may be damaged. The RPM sensor is then fastened to the motor
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hub of the model such that the marked surface of the rotor is approximately 1 mm away
from the tip of the sensor. The signal of the RPM sensor can be monitored through a
digital spectrum analyzer via a BNC cable spliced into the sensor wire and through a
laptop computer via USB. The USB cable is plugged into the “USB” port of the elogger.
Figure 6-30: 10 series model electronics setup
For each of the three models tested in the wind tunnel, the same rotor was used.
The selected rotor was the MA 9.5”x 5” which has 3 blades. Also, only the smallest tip
clearance of 1%R was used during the experiments. Three model angles of attack were
used throughout the wind tunnel experiments including -3o, 0 o, and 3 o. These angles
were selected since it was anticipated that a vehicle utilizing ducted fans for lift would
operate in forward flight at shallow angles of attack. For models 10-1 and 10-2, the angle
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of attack adjustment bar was attached directly to the models. The angle of attack was
checked by using a digital level across the exit of the duct. As shown in Figure 6-31, all
of the wires external to the model exited the test section along the angle of attack
adjustment bar. The wires were also covered and fastened to the model using the HVAC
tape. Figure 6-31 also illustrates how the holes in the floor of the test section were sealed
with the tape. A foam backing was applied under the tape on the floor to provide extra
rigidity to the seal.
Figure 6-31: Wire routing to angle of attack adjustment bar for 10 series ducted fan model inside APB wind tunnel
Model 10-2 was tested in the wind tunnel with the vents opened and closed.
Figure 6-32 shows the model inside the test section with the vents open. The HVAC tape
was applied to both the inside and outside of the duct when it was desired to test the
model with the vents closed. Figure 6-33 illustrates the tape placement for the closed
vent configuration. 188
Figure 6-32: Open vent configuration of model 10-2 inside the APB wind tunnel
Figure 6-33: Closed vent configuration of model 10-2 inside the APB wind tunnel
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The angle of attack adjustment bar was not connected to the isolated rotor model.
The angle of attack was adjusted by pivoting the model at the connection points to the
force balance struts. The angle of attack was then checked by placing a digital level
across the bottom of the motor mount. As shown in Figure 6-34, the wires connected to
the isolated rotor model were run along the threaded rod of the model and down one of
the force balance struts. The HVAC tape was used again to secure the wire to the
threaded rod and strut.
Figure 6-34: Isolated rotor model mounted inside the APB wind tunnel
6.6 Experiment Procedure
Before any wind tunnel test could begin, the three access hatches were removed between the first and second corners of the wind tunnel. The wire mesh net just upstream
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of the first corner was inspected for any loose particles that may be present from previous
experiment. Also, the turning vanes in the first corner were inspected through the first
access hatch. The second access hatch is located centrally between the first and second
corners. The guide vanes in the second corner as well as the fan drive shaft were
inspected through this hatch. Also, the interior of the tunnel was inspected at this station.
The third access hatch allows for the wind tunnel fan itself to be inspected. Any potential
issues such as loose bolts or leaking oil can be seen at that location. Proper attention
should be given when inspecting the tunnel through these hatches. Power to the tunnel
should especially not be turned on when inspecting the tunnel fan.
Once the tunnel inspection is completed and there are no signs of potential danger, the power to the tunnel can be applied. Typically, the tunnel motor cooling fan is first powered on simply by flipping the switch to the circuit breaker box. Next, the power to the AC inverter drive can be applied, again by simply turning the handle on the exterior of the unit. Once this handle is turned the wind tunnel motor will be ready to use. The device used to control the power supplied to the wind tunnel motor, and hence wind tunnel speed, is located at the data acquisition station platform. The small control is hard-wired to the AC inverter drive which is located at the rear of the wind tunnel facility room. The hand-held control features basic buttons such as a start and stop button as well as increasing or decreasing input frequency buttons. Recall that the wind tunnel speed is controlled by the AC inverter drive. The control is setup up so that the frequency of the input power to the tunnel motor can be varied. The tunnel speed was calibrated with the motor input frequency. The results of this calibration are presented in Figure 6-35.
Therefore, the wind tunnel speed can be obtained for a given input power frequency.
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Once the particular frequency is dialed and the start button is pressed, the tunnel motor will gradually increase its RPM to reach the desired speed. The tunnel speed can also be simultaneously adjusted while the motor is running. If the stop button is pressed, the tunnel motor will gradually decrease its RPM until a zero velocity is obtained in the test section. Before a velocity is applied to the test section however, a tare procedure must be done to the voltage channels of the force balance and pressure difference in the contraction section.
80
70 y = 4.3953x 60 R2 = 0.9963 50
40
30
20
Wind Tunnel Velocity [ft/sec] 10
0 0 5 10 15 20 Motor Input Frequency [Hz]
Figure 6-35: Wind tunnel motor calibration
This tare correction procedure is identical for all the models tested. When the experiment is ready to begin with the model set at a desired angle of attack, the force and moment balance should be unlocked. Then the auto-tare switches under each channel of
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the balance strain amplifier should be pressed as similarly done during the hover thrust
experiments. When the wind tunnel LabView code is running, the instantaneous voltages
of the balance will go to zero. The manual tare knobs can be used for finer adjustment.
Next, the voltage for the pressure difference in the contraction section should be zeroed using the tunnel amplifier as described earlier, either using the zero knob or zeroing function in the LabView program. Once both of these were zeroed, the first set of data was acquired for a zero rotor RPM and zero wind tunnel speed case using the LabView program. This was the true zero load case. It is after this first acquisition that the desired velocity is applied to the test section. Once the tunnel motor reaches a steady state test section velocity, a second zero rotor RPM data set is acquired. Since the rotor is not turning, this point essentially corresponds to the aerodynamic performance of just the duct or support structure. With the two zero loading cases acquired, the rotor RPM was then stepped to a maximum value using the servo tester. A data point was acquired at each RPM increment. Once the data was acquired for the maximum rotor RPM, the two zero loading cases were then repeated in reverse order and the stop button on the tunnel control was pressed. The sequence just described represents one trial for a fixed wind tunnel speed and a fixed angle of attack. For repeatability purposes, this identical procedure was repeated for a total of three trials at any given wind tunnel speed and model angle of attack for models 10-1 and 10-2. Only two trials were completed for the isolated rotor model. When a higher wind tunnel speed was desired, a different tunnel input power frequency was simply dialed into the control. The angle of attack was also set at the beginning of the first trial.
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The wind tunnel experiments were planned such that the various models would be tested over a wide range of advance ratios. The lower range was to simulate the transition from hover to forward flight. The performance of the models at high speed flight was also tested at high values of advance ratios. Additionally, many cruising speeds between hover transition and dash speeds were tested. The forward flight test matrix for each model is presented in Tables 6-1, 6-2, and 6-3. The initial model input voltage was set to 10 V with the DC power supply. The input current was allowed to adjust to the required amount at any given rotor RPM.
Table 6-1: Model 10-1 forward flight test matrix
Duct AoA Rotor Installed ? Wind Tunnel Speed Rotor RPM Advance Ratio Range [Deg] [ft/sec] [RPM] [ ] 10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 YES 50 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.2 - 0.48 62 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.25 - 0.6 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64 -3 45 50 NO 56 62 67 70
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 YES 50 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.2 - 0.48 62 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.25 - 0.6 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64 0 45 50 NO 56 62 67 70
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 YES 50 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.2 - 0.48 62 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.25 - 0.6 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64 3 45 50 NO 56 62 67 70 Notes: 1. Input power supply voltage was set at 10 V 2. 3 trials of each case were performed for repeatability and averaging 3. A single tip clearance of 1% R was used with the MA 10x5 rotor
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Table 6-2: Model 10-2 forward flight test matrix
Duct AoA Rotor Installed Wind Tunnel Speed Rotor RPM Advance Ratio Range [Deg] YES/NO [ft/sec] [RPM] [ ] 10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 YES 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64
10 -3 25 45 NO 50 56 62 67 70
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 YES 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64
10 0 25 45 NO 50 56 62 67 70
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 YES 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64
10 3 25 45 NO 50 56 62 67 70 Notes: 1. Input power supply voltage was set at 10 V 2. 3 trials of each case were performed for repeatability and averaging 3. A single tip clearance of 1% R was used with the MA 10x5 rotor 4. This matrix was completed with the vents opened and closed
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Table 6-3: Isolated rotor forward flight test matrix
TPP AoA Wind Tunnel Speed Rotor RPM Advance Ratio Range [Deg] [ft/sec] [RPM] [ ] 10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 -3 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 0 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64
10 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.04 - 0.1 3 25 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.1 - 0.24 66 2500, 3000, 3500, 4000, 4500, 5000, 5500, 6000 0.27 - 0.64 Notes: 1. TPP = Tip Path Plane 2. Input power supply voltage was set at 10 V 3. 2 trials of each case were performed for repeatability and averaging 4. A single tip clearance of 1% R was used with the MA 10x5 rotor
6.7 Forward Flight Experiment Results
During the forward flight experiments, seven model configurations were tested in the APB wind tunnel. These included an isolated rotor as well as ducted fan models 10-1 and 10-2. Recall that the rotor used in these experiments was the MA 9.5”, 3 blade, rotor with square tips. For both models 10-1 and 10-2, the rotor tip clearance was kept at 1%
RR. In the specific case of model 10-2, measurements were taken for the forward inlet vents opened and closed. Finally, measurements were also taken for the 10-1 and 10-2 ducts only (without rotor). This section presents the highlights from these experiments.
Figure 6-36 presents the measured dimensional lift as a function of wind tunnel speed for the isolated rotor (MA 9.5”), ducted rotor, and isolated duct, all at a 0 o angle of attack. Thus each model is subjected to an edgewise flow. In the isolated rotor and ducted rotor cases, the rotor RPM was kept at a constant 6000 RPM while the wind
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tunnel speed was varied. Notice that for each model, the lift increases with wind tunnel
speed. Of interest is the difference in lift produced by the isolated rotor and ducted rotor.
For the same operating conditions, each ducted rotor produced nearly three times the
amount of lift compared to the isolated rotor.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, Isolated Rotor, 6000 RPM 1% R TC 10-1 (6000 RPM) 2 10-2 Vents Closed (6000 RPM) 10-2 Vents Open (6000 RPM) 10-1 Duct Only 10-2 Duct Only (vents closed) 1.5 10-2 Duct Only (vents open)
1 Lift [lb]
0.5
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-36: Dimensional lift as a function of wind tunnel speed for the isolated rotor, isolated duct, and ducted fan models
Comparing just the ducted rotor models at 6000 RPM, models 10-1 and 10-2 with the vents closed produced about the same amount of lift over all wind tunnel speeds except for the highest speed tested (70 ft/sec). At that speed, model 10-2 with the vents closed produced about 10% more lift than 10-1. Model 10-2 with the vents open produced the lowest amount of lift at speeds lower than 40 ft/sec, but it produced the most amount of lift compared to the other ducted rotor models at 70 ft/sec.
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Finally Figure 6-36 shows that the ducts by themselves produce little lift, even at
high wind tunnel speeds. It is interesting to note that the model 10-2 duct, with both the
vents open and closed, produced more lift than the model 10-1 duct. However, the
highest amount of duct lift was only about 30% of the lift produced by the isolated rotor.
Thus, the ducts alone produced little body lift. This was to be expected since the single
duct is a fairly blunt object with little lifting surface. A vehicle with integrated ducted lift
fans would need to get body lift from the external structure surrounding the duct or ducts.
The lift data just shown for the ducted rotor models was then non-dimensionalized with wind tunnel speed. Figure 6-37 presents the lift coefficient as a function of wind tunnel speed for models 10-1 and 10-2 (vents closed and opened) operating at 6000 RPM.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 10 10-1 (6000 RPM) 9 10-2 vents closed (6000 RPM) 8 10-2 vents open (6000 RPM) 7 6 WT )
L 5 (C 4 3 2 1 0 0 20 40 60 80
VWT [ft/sec]
Figure 6-37: Lift (non-dimensionalized with wind tunnel speed) as a f unction of wind tunnel speed for each ducted fan model at 6000 RPM
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Up to wind tunnel speeds of 25 ft/sec, model 10-1 appears to produce the largest lift
coefficient. A maximum lift coefficient of almost 10 is obtained. Meanwhile, model 10-
2 with the vents open produced the lowest lift coefficient in that speed range. This is
consistent with the previous dimensional data. However, the lift coefficient behavior of
each model with the rotor operating at 6000 RPM follows a different trend. At low
speeds the lift coefficient is large but as the wind tunnel speed increases, the lift
coefficient of each model decreases. Thus, when the lift is non-dimensionalized in this
manner, the lift coefficient is seen to decrease with increasing wind tunnel speed.
Figure 6-38 presents the same lift coefficient (non-dimensionalized with wind tunnel speed) as a function of advance ratio for each ducted fan model. The advance ratio essentially non-dimensionalizes wind tunnel speed with rotor tip speed. For this test, the wind tunnel speed varied from 10 ft/sec to 70 ft/sec. The rotor RPM was also varied from 2500 to 6000 RPM. Essentially, low advance ratios correspond to low wind tunnel speed and high rotor RPM. High advance ratios correspond to high wind tunnel speed and low rotor RPM. Furthermore, the typical range of advance ratios for modern helicopters is 0 to 0.5, (0.5 being high speed forward flight). Thus, low advance ratios correspond to the transitional stage in the flight envelope from hover to low speed forward flight.
Figure 6-38 shows that the lift coefficient of each ducted fan model, fixed at 0 o angle of attack, decreases with increasing advance ratio. This trend is consistent with
Figure 6-37. Advance ratios in the range of 0 to 0.1 represent the transition from hover to low speed forward flight. Figure 6-38 suggests that when the ducted fan is fixed at a 0 o
angle of attack, the lift coefficient is maximized in this transitional flight advance ratio
199
range. Upon comparison to the other ducted fan models in this range, model 10-2 with
the vents open produced the lowest lift coefficient.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 10 10-1 9 10-2 vents closed 8 10-2 vents open 7 6 WT )
L 5 (C 4 3 2 1 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio [V WT /V Tip ]
Figure 6-38: Lift (non-dimensionalized with wind tunnel speed) as a function advance ratio for each ducted fan model, V WT and V Tip variable
The lift data for each ducted fan with the MA 9.5” rotor operating at 6000 RPM
was also non-dimensionalized with rotor tip speed. Figure 6-39 shows the lift coefficient
(non-dimensionalized with rotor tip speed) as a function of wind tunnel speed for each
ducted fan model at a 0 o angle of attack. In general, the lift coefficient increases with wind tunnel speed. This contradicts the previous trend when lift was non- dimensionalized with wind tunnel speed. Also in contrast, Figure 6-39 shows that model
10-1 produced the highest lift coefficient for all wind tunnel speeds. Model 10-2 with the vents open produced the lowest lift coefficient compared to the other models up to a
200
speed of 50 ft/sec. After 50 ft/sec, model 10-2 with the vents open produced a higher lift
coefficient that of model 10-2 with the vents closed but still lower than model 10-1.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.025
0.02
0.015 Tip ) L (C 0.01
10-1 (6000 RPM) 0.005 10-2 vents closed (6000 RPM) 10-2 vents open (6000 RPM) 0 0 20 40 60 80
VWT [ft/sec]
Figure 6-39: Lift (non-dimensionalized with rotor tip speed) as a function of wind tunnel speed for each ducted fan model at 6000 RPM
The lift coefficient, non-dimensionalized with rotor tip speed, for each ducted fan
model at 0 o angle of attack was also plotted as a function of advance ratio. As shown in
Figure 6-40, the wind tunnel speed and rotor RPM were varied. For all the models, the
lift coefficient is shown to increase with advance ratio. At advance ratios less than 0.3,
model 10-1 produced the highest lift coefficient and model 10-2 with the vents open produced the lowest lift coefficient. After advance ratios of 0.3, model 10-2 in both
configurations produced higher lift coefficients than model 10-1. It also appears that at
advance ratios below 0.1, the lift coefficient is minimized. Recall that a 0 advance ratio
201
represents the hover condition. Referring to Figure 5-25 from Chapter 5, the thrust in
hover was also non-dimensionalized with rotor tip speed. In hover, the maximum thrust
coefficient for model 10-1 with 2%R tip clearance was 0.027. Thus, Figure 6-40 shows
that as the single ducted fan transitions from hover to forward flight, the lift coefficient
decreases to a minimum at low forward speeds and then increases as the forward speed becomes greater.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.035
0.03
0.025
0.02
Tip T ) L o (C 0.015 H o 0.01 v e 10-1 0.005 r 10-2 vents closed 10-2 vents open 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio [V WT /V Tip ]
Figure 6-40: Lift (non-dimensionalized with rotor tip speed) as a function of advance ratio for each ducted fan model, V WT and V Tip variable
Next, the drag was examined compared to the lift. Figure 6-41 presents the dimensional lift and drag produced by each of the ducted fan models using the MA 9.5” rotor at 6000 RPM. Just as lift, the drag also increases with wind tunnel speed. At wind tunnel speeds less than about 30 ft/sec, each ducted fan model produced more lift than
202
drag. However, as the wind tunnel speed increased above 30 ft/sec, the drag produced by
each ducted fan model became greater that the lift. At the maximum speed tested of 70
ft/sec, model 10-1 produced a drag approximately equal to two times the amount lift it produced. This suggests that a single ducted fan in edgewise flight has poor aerodynamic efficiency at high forward speeds.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC, 6000 RPM 3.5 (Drag) 10-1 (Drag) 10-2 vents closed 3 (Drag) 10-2 vents open (Lift) 10-1 2.5 (Lift) 10-2 vents closed (Lift) 10-2 vents open 2
1.5
1 Lift [lb] and Drag [lb]
0.5
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-41: Dimensional lift and drag as a function of w ind tunnel speed for each ducted fan model at 6000 RPM
The measured drag was further examined in Figure 6-42. The dimensional drag was plotted as a function of wind tunnel speed for each isolated duct at an angle of attack of 0 o. Recall that in the duct only configuration, the rotor was removed from the duct.
For all of the ducts, drag increased with wind tunnel speed. The model 10-2 duct with
the vents open produced the largest amount of drag, especially at high wind tunnel
203 speeds. At the maximum speeds tested, the model 10-2 duct with vents opened produced approximately two times the drag of both of the other ducts.
0 Deg AoA 3.5 10-1 Duct Only 3 10-2 Duct Only (vents closed) 10-2 Duct Only (vents open) 2.5
2
1.5 Drag [lb]
1
0.5
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-42: Comparison of dimensional drag for each isolated duct at 0 o angle of attack
Figure 6-43 presents the dimensional drag plotted as a function of wind tunnel speed for each of the ducted fan models as well as the isolated rotor operating at 6000
RPM and 0 o angle of attack. First notice that the drag produced by the isolated rotor is nearly constant over all wind tunnel speeds. The drag shown for the isolated rotor was corrected for the tare drag from the rotor support structure and force balance struts. Next, notice that the drag produced by each ducted fan model at 6000 RPM increases steadily with wind tunnel speed. Thus, the ducted fan models produce substantially more drag than that of the isolated rotor. Comparing each ducted fan model, 10-2 with the vents
204
open produced the lowest drag at low wind tunnel speeds. However, at the highest wind
tunnel speed model 10-2 with the vents open produced the highest drag coefficient.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 3.5 Isolated Rotor, 6000 RPM 3 10-1 (6000 RPM) 10-2 Vents Closed (6000 RPM) 10-2 Vents Open (6000 RPM) 2.5
2
1.5 Drag [lb]
1
0.5
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-43: Comparison of dimensional drag for isolated rotor and each ducted fan model at 6000 RPM
Upon comparison to the drag produced by the ducts alone of Figure 6-42, it
appears that the drag measured for each ducted fan operating at 6000 RPM is not equal to just the duct alone and isolated rotor operating at 6000 RPM. The drag measured for each ducted fan as shown in Figure 6-43 is actually greater than just the combination of isolated duct and rotor. This extra component of measured drag is momentum drag.
Recall that a result of the rotor forcing the free-stream velocity to turn vertically through the axis of the duct is the production of momentum drag. Horn[7] states that momentum
205 drag is a function of forward speed, mass flow rate through the rotor, and duct turning efficiency.
The drag force produced by each ducted fan model operating at 6000 RPM was then non-dimensionalized with wind tunnel speed. Figure 6-44 shows the drag coefficient (corrected for tare drag) as a function of wind tunnel speed for each ducted fan model at a 0 o angle of attack. For each model, the drag coefficient is maximized at low wind tunnel speeds and decreases with increasing wind tunnel speed. Thus, even though the dimensional drag increases, the drag coefficient decreases with forward speed.
Interestingly, model 10-2 with the vents open produced the lowest drag coefficient up to wind tunnel speeds of about 55 ft/sec.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 3.5 10-1 (6000 RPM) 10-2 vents closed (6000 RPM) 3 10-2 vents open (6000 RPM)
2.5
2 corrected )
D 1.5 (C
1
0.5
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-44: Drag coefficient, corrected for tare drag, as a function of wind tunnel speed for each ducted fan model at 6000 RPM
206
Figure 6-45 shows the drag coefficient of each ducted fan model at a 0 o angle of
attack as a function of advance ratio. For advance ratios of less than 0.2, the drag
coefficient of model 10-2 with the vents open appears to be the lowest of all the ducted
fan configurations. At an advance ratio of 0.05, the drag coefficient for 10-2 with the
vents open is 30% less than that with the vents closed. However, at advance ratios
greater than 0.2, the drag coefficient reduction is not sustained by the open vent
configuration and it increases. Therefore, for drag coefficient considerations, it would
appear that the open vent configuration could be useful during hover transition to low-
speed forward flight. As the vehicle establishes a steady forward flight, the vents could
then close in order to avoid the increase in drag coefficient.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 3.5 10-1 3 10-2 vents closed 10-2 vents open 2.5
2 corrected )
D 1.5 (C
1
0.5
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio [V WT /V Tip ]
Figure 6-45: Drag coefficient as a function of advance ratio for each ducted fan model, VWT and V Tip variable
207
The pitching moment produced by each model was also examined. Figure 6-46 presented the dimensional pitching moment of each isolated duct as a function of wind
tunnel speed. The models were mounted to the force and moment balance such that the pitching moment was measured about the 3/4 duct chord location. Each isolated duct
model produced a nose-up pitching moment that increases with forward speed.
0 Deg AoA 1.2 10-1 Duct Only 10-2 Duct Only (vents closed) 1 10-2 Duct Only (vents open)
0.8
0.6
0.4 Pitching Moment [lb ft] 0.2
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-46: Dimensional pitching moment about 3/4 chord of each isolated duct at 0 o angle of attack as a function of wind tunnel speed
Fleming[13] suggests that a cause for the nose-up pitching moment of ducted fans is that
the center of pressure that the drag acts through is actually outside and above the duct,
thus creating a moment arm. Another possible explanation for the nose-up pitching
moment, suggested by Horn[7], is that the duct lip experiences asymmetric lift in forward
flight. The leading edge inlet lip produces a greater amount of lift than the trailing edge
lip, hence creating the moment. Comparing each of the isolated ducts, the model 10-2
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duct with vents open produced the highest pitching moment over the entire range of wind
tunnel speeds. The closed vent configuration of the model 10-2 duct produced the lowest pitching moments over all wind tunnel speeds.
Figure 6-47 shows the dimensional pitching moment (about 3/4 duct chord) for
each ducted fan model and isolated rotor operating at 6000 RPM. Both the isolated rotor
and ducted fan models were shown to produce increasing nose-up pitching moments with
forward speed. Interestingly, model 10-2 with the vents open produced the lowest pitching moment over the entire range of wind tunnel speeds when the rotor was
operating at 6000 RPM. This is the opposite as expected based on the pitching moment produced by the isolated model 10-2 duct with vents open.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 1.6 Isolated Rotor, 6000 RPM 1.4 10-1 (6000 RPM) 10-2 Vents Closed (6000 RPM) 1.2 10-2 Vents Open (6000 RPM)
1
0.8
0.6
Pitching Moment [lb ft] 0.4
0.2
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-47: Dimensional pitching moment for each ducted fan model and isolated rotor at 6000 RPM as a function of wind tunnel
209
The pitching moment was also non-dimensionalized with wind tunnel speed.
Figure 6-48 shows the pitching moment coefficient as a function of wind tunnel speed for each ducted fan model operating a 6000 RPM and 0 o angle of attack. It was found that each ducted fan produced a large nose-up pitching moment coefficient at low wind tunnel speeds. As the forward speed increased, the pitching moment coefficient was reduced.
Upon comparison of each ducted fan, model 10-1 produced the largest pitching moment coefficient. Model 10-2 with the vents open produced the lowest pitching moment coefficient over the entire range of wind tunnel speeds. The pitching moment coefficient for the open vent configuration is actually reduced by a factor of 2 when compared to 10-
1 at low speeds.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 6 10-1 (6000 RPM) 10-2 vents closed (6000 RPM) 5 10-2 vents open (6000 RPM)
4 WT )
m 3 (C
2
1
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-48: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) for each ducted fan model at 6000 RPM as a function of wind tunnel speed
210
Figure 6-49 presents the pitching moment coefficient as a function of advance
ratio for each ducted fan model at a 0 o angle of attack. Most significant was the reduction
of the pitching moment coefficient produced by model 10-2 with the vents open for
advance ratios less than 0.2. At an advance ratio of 0.05, the open vent configuration
actually produced a pitching moment coefficient of almost 35% less than that of the
closed vent configuration. After advance ratios of 0.2, however it appears that the pitching moment coefficient of model 10-2 with vents open begins to increase over that
of the closed vent case.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 6 10-1 5 10-2 vents closed 10-2 vents open
4 WT )
m 3 (C
2
1
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio [V WT /V Tip ]
Figure 6-49: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) for each ducted fan model as a function of advance ratio, V WT and V Tip variable
When the pitching moment was non-dimensionalized with rotor tip speed, the pitching moment coefficient was found to increase with wind tunnel speed as shown in
211
Figure 6-50. For each ducted fan operating at 6000 RPM and 0 o angle of attack, an increasing positive pitching moment coefficient was produced. The data shows that model 10-2 with the vents open again produced the lowest pitching moment coefficient over the entire speed range while model 10-1 produced the highest.
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.06 10-1 (6000 RPM) 10-2 vents closed (6000 RPM) 0.05 10-2 vents open (6000 RPM)
0.04 Tip )
m 0.03 (C
0.02
0.01
0 0 20 40 60 80
VWT [ft/sec]
Figure 6-50: Pitching moment coefficient (non-dimensionalized with rotor tip speed) for each ducted fan model at 6000 RPM as a function of wind tunnel speed
Figure 6-51 presents the pitching moment coefficient non-dimensionalized with rotor tip speed as a function of advance ratio for each ducted fan model at 0 o angle of
attack. It appears that the pitching moment coefficient produced by model 10-2 with the
vents open is the lowest compared to the other models for advance ratios less than 0.3.
212
0 Deg AoA MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.2 10-1 0.18 10-2 vents closed 0.16 10-2 vents open
0.14
0.12 Tip )
m 0.1 (C 0.08
0.06
0.04 0.02
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio [V WT /V Tip ]
Figure 6-51: Pitching moment coefficient (non-dimensionalized with rotor tip speed) for each ducted fan model as a function of advance ratio, V WT and V Tip variable
So far, all of the forward flight data presented has been for a fixed angle of attack of 0 o. The final series of plots present the non-dimensional forces and moments as a
function of the advance ratio and angle of attack for the 10-1 model only. Model 10-1
was used as a representative case since the trends are similar for all the duct
configurations tested. Figure 6-52 presents the drag coefficient versus advance ratio for
model 10-1 at three angles of attack. Notice that for advance ratios less than 0.1, the drag
coefficient data diverges and do not collapse to a single curve. This is because a
component of the rotor thrust vector was being measured along with the drag through the
drag channel of the force balance. Recall that the 10-1 model was mounted inverted to
the balance inside the test section. Thus for a positive 3 o angle of attack, a component of
213 the rotor thrust vector was pointing in the positive drag direction. Essentially, the rotor thrust was being added to the drag being measured. The drag was measured to be less for the -3o angle of attack since a component of the rotor thrust was being directed opposite the drag direction. At advance ratios greater than 0.1, the drag coefficient diverges slightly, and the drag of the model becomes greater than the horizontal component of the rotor thrust.
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 4
3.5 -3 Deg AoA 0 Deg AoA 3 +3 Deg AoA
2.5
2 corrected ) D
(C 1.5
1
0.5
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-52: Drag coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
Figure 6-53 shows the side-force coefficient as a function of advance ratio. As seen, the side-force coefficient on average is an order of magnitude less than the drag coefficient. Also, there is enough scatter within the data such that a difference between
214 angle of attack cases is not observable. Generally however, there is a positive, port side, side-force over all advance ratios.
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.25 -3 Deg AoA 0 Deg AoA 0.2 +3 Deg AoA
0.15 WT ) S (C 0.1
0.05
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-53: Side force coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
Figure 6-54 presents the lift coefficient (non-dimensionalized with wind tunnel speed) as a function of advance ratio for the 10-1 ducted fan. It does not appear that the duct angle of attack has an effect on the lift coefficient throughout all advance ratios.
This trend was also observed for differently shaped ducts that were tested. It should be noted that this was observed over a shallow angle of attack change. Although the edgewise ducted lift fan would likely only be subjected to shallow angles of attack when integrated into a complete vehicle, it would be interesting to see how larger angles of attack affect the lift coefficient. The lift was also non-dimensionalized with rotor tip speed and is shown in Figure 6-55 as a function of advance ratio.
215
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 12 -3 Deg AoA 0 Deg AoA 10 +3 Deg AoA
8 WT )
L 6 (C
4
2
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-54: Lift coefficient (non-dimensionalized with wind tunnel speed) of model 10- 1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.04
0.035
0.03
0.025 Tip )
L 0.02 (C 0.015
0.01 -3 Deg AoA 0.005 0 Deg AoA +3 Deg AoA 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-55: Lift coefficient (non-dimensionalized with rotor tip speed) of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
216
With the lift non-dimensionalized with rotor tip speed, there does appear to be a change in lift coefficient with angle of attack at high advance ratios as shown in Figure 6-
55. At advance ratios greater than 0.4, it appears that the lift coefficient increases with angle of attack.
Figure 6-56 shows the rolling moment coefficient for model 10-1 as a function of advance ratio. The magnitude of the rolling moment coefficient is small compared to the other moments measured. Also after advance ratios of about 0.05, the rolling moment coefficient remains positively constant over all advance ratios and duct angles of attack.
This corresponds to a constant port-side rolling moment. However, at advance ratios of less then 0.05, there is a rolling moment coefficient of opposite sign.
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.2 0.15 0.1 0.05 0 0.05 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 WT ) l
(C 0.1 0.15 0.2 0.25 -3 Deg AoA 0 Deg AoA 0.3 +3 Deg AoA 0.35
Advance Ratio, V WT /V tip
Figure 6-56: Rolling moment coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
217
The ducted fan essentially experiences a starboard rolling moment at low advance
ratios, and then reverses direction to a port-side rolling moment. Thus for the single fan
model 10-1, there exists an advance ratio where a trim rolling moment condition is
obtained. In this case, that trim condition advance ratio is 0.05. Although the magnitude
of the rolling moment coefficient may be low, the change in direction could cause further performance concerns during the low advance ratio or transitional flight mode.
The torque of the motor was measured with the yaw moment channel of the force balance. The yaw moment coefficient as a function of advance ratio and duct angle of attack is shown in Figure 6-57. This is the duct’s reaction to the rotational torque of the rotor. Notice that the yawing moment coefficient decreases with increasing advance
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 1.2 -3 Deg AoA 0 Deg AoA 1 +3 Deg AoA
0.8 WT )
n 0.6 (C
0.4
0.2
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-57: Yawing moment coefficient of model 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
218 ratio. Recall that the advance ratio is a function of the inverse of rotor tip speed. Thus for low advance ratios, the rotor tip speed is high. For high advance ratios, the rotor tip speed is low. Therefore, the measured yawing moment should indeed be highest for lowadvance ratios as is shown. There also appears to be no dependence on duct angle of attack.
Figure 6-58 presents the pitching moment coefficient (non-dimensionalized with wind tunnel speed) as a function of advance ratio and duct angle of attack for model 10-1.
Using this method of non-dimensionalization yields no dependence on angle of attack.
However, when the pitching moment is non-dimensionalized with rotor tip speed, as shown in Figure 6-59, the pitching moment coefficient appears to change with angle of attack. At advance ratios greater than 0.3, the pitching moment coefficient appears to decrease with increasing duct angle of attack.
219
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 7 -3 Deg AoA 6 0 Deg AoA +3 Deg AoA 5
4 WT ) m
(C 3
2
1
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-58: Pitching moment coefficient (non-dimensionalized with wind tunnel speed) of 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
Model 10 1 MA 9.5", Square Tip, 3 blade rotor, 1%R TC 0.25 -3 Deg AoA 0 Deg AoA 0.2 +3 Deg AoA
0.15 Tip ) m
(C 0.1
0.05
0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7
Advance Ratio, V WT /V tip
Figure 6-59: Pitching moment coefficient (non-dimensionalized with rotor tip speed) of 10-1 ducted fan as a function of advance ratio and angle of attack, V WT and V Tip variable
220
Chapter 7
Remarks and Conclusions
The work presented in this thesis has attempted to uncover the basic and
fundamental phenomena associated with ducted fan technologies. Of particular interest
was the subjecting of an edgewise flow to a ducted fan. If integrated into a total air
vehicle system, this type of ducted fan would be responsible for providing the majority of
the lift. Some of the questions that were answered included how the duct changes the
velocity profile of the rotor, how much thrust augmentation the duct provides, and how
the rotor tip clearance affects the hover thrust performance. The performance of single
ducted fans in forward flight was also examined. The lift, drag, and pitching moment behavior was determined for a single ducted fan throughout a range of forward flight
conditions, including the transitional stage between hover and low-speed forward flight.
Also, the constraints of ducted lift fans in edgewise flight were highlighted and a method
for potentially reducing these limitations was introduced. This project was approached
through a design, build, and test mentality. The next section briefly summarizes each of
the preceding chapters and highlights the most significant points of the three-step
approach. The chapter will close with suggestions for future work to be conducted in this
area.
221
7.1 Summarizing Remarks
Chapter 1 served to introduce the ducted fan concept. Potential advantages of the
use of ducted fans were identified such as the ability to fly in close proximity to
obstacles, less complicated control hub structures, and less noise propagation than
traditional open rotor systems. Also, the idea that the duct allows for an increase in
overall thrust meant that a ducted fan could produce the same amount of thrust as a larger
open rotor. Some of the more pertinent geometric characteristics of the ducted fan were
also introduced including inlet lip radius, diffuser exit angle, and rotor tip clearance. The
rotor tip clearance was defined as the gap distance between the rotor tip and the inside
wall of the duct.
Next, an overview of the historical uses of ducted fans in air vehicles was given beginning in the 1940’s up until the current time. This historical look was meant to provide a glimpse of the evolution of ducted fan technology, where it succeeded, and where it failed. Several examples were provided including the use of ducted fans in fixed wing propulsion, rotary wing propulsion, tilt duct aircraft, and direct lift VTOL vehicles.
The historical review showed how current state of the art ducted fan vehicles bare close resemblance to the ideas of the past. For example, the Joint Strike Fighter is very similar to the Ryan XV-5. The Urban Aeronautics X-hawk closely resembles the Piasecki Air
Jeeps. Finally, an important distinction was made that would set the tone for the rest of the thesis. It was proposed that all ducted fan air vehicles could be categorized into two groups. There are those that use ducted fans for providing the thrust of the vehicle and those that use the ducted fans solely for providing lift. (There are also air vehicles that
222
utilize ducted fans for both purposes). Thus the focus the current research efforts was placed on ducted fans designed to provide lift.
Chapter 2 presented an aerodynamic model for open rotors and ducted rotors, both in hover and in forward (edgewise) flight. This was done by treating the rotor in
each case as an actuator disk and using momentum theory. It was assumed that for an
ideal ducted rotor in hover, the duct itself produces an amount of thrust equal to that of
the rotor. Upon comparison of an ideal ducted rotor in hover producing the same amount
of thrust as an open rotor of equal disk area, it was found that ducted rotors can achieve a
30% savings in power. In other words, an ideal ducted rotor requires less power to produce the same amount of thrust as that of an open rotor of equal size. An alternative
conclusion was also shown such that an ideal ducted rotor can produce the same amount
of thrust for the same power required as an open rotor of twice the disk area. Thus an
ideal ducted rotor can be smaller than an open rotor without sacrificing thrust.
Two methods were also presented to account for non-ideal conditions of the
ducted rotor in hover such as non-uniform inflow. One method introduced a thrust
augmentation factor. This allowed for the duct to produce a thrust equal to a fraction of
the rotor thrust. When the limits of this thrust augmentation factor were defined properly,
it was shown that the same equations for an ideal ducted rotor as well as an open rotor
could be backed out. The second non-ideal ducted rotor method used a wake contraction parameter which made the far wake area a function of the rotor disk area. Again, with proper limits of this parameter, the previous equations were able to be retrieved.
The momentum theory was then applied to open rotors and non-ideal ducted
rotors in forward edgewise flight. Some important factors introduced for the non-ideal
223
ducted rotor in forward flight were the turning efficiency factor and momentum drag.
The turning efficiency factor describes how well the duct turns the air flow from
horizontal to vertical through the duct. Momentum drag is produced as the air is forced
to suddenly change direction through the duct. Thus, a high turning efficiency leads to
higher momentum drag. This analysis revealed tradeoffs that are frequently present in
aircraft design.
With an understanding of past designs and results from a basic momentum theory
analysis, Chapter 3 details several ducted fan models, built in house, that would be used
for the purposes of performing experiments. A general design methodology was laid out
including trade studies that were performed comparing ducted fans used in government
and industry research. Specific geometries were non-dimensionalized for comparison
such as duct chord and duct thickness. Rotor characteristics were also studied in a similar
fashion. Several models were fabricated, ranging in sizes from 15 inches to 10 inches in
diameter.
The first model built (the Ford Fan model), featured a 15.25 inch diameter, 10 bladed radiator fan surrounded by a foam duct. This model was the largest of all the
models built and served as an initial exercise in fabrication and experiment techniques.
Next came the LL1 series models. These models used 14.5 inch diameter, 4 bladed
rotors. The LL1 models featured additional capabilities over the previous model through
upgraded motors, varying duct exit angles, and varying rotor tip clearance cases. The
final sets of models built were titled the 10 series models. These models were designed
with wind tunnel testing in mind. Thus they were nominally 10 inches in diameter.
Advancements to the fabrication of these models included fiberglass composite duct
224 structures for extra strength. In general, the 10 series models kept the same capabilities as the previous models with a few additional enhancements. Model 10-2 incorporates the venting of the forward portion of a ducted fan, as seen in the Bell Helicopter/Urban
Aeronautics case. In addition to a smaller 4 bladed rotor, a 3 bladed rotor was also used in the 10 series models. Thus variations in rotor performance could be examined with the use of rotors with a different blade number, thickness, and pitch. One of the other important additions to the 10 series models was instrumentation to better control the rotor
RPM and monitor motor temperature. With all the models, the ducts could also be removed or a separate setup could be used to investigate the performance of an isolated rotor. Overall, the capability was developed to produce high-quality models that are well suited for experiments over a wide range of parameters.
Chapter 4 presented the first series of aerodynamic experiments with the newly fabricated ducted fan models. The velocity profiles of two different ducted fan models were measured. The proper facilities were matured while keeping safety a priority.
The velocity distribution along the blade span was found for the LL1 models.
Two 4-bladed rotors of varying geometry were studied. One of the rotors also had tapered blade tips, whereas the other had square blade tips. It was shown that both isolated rotors achieved a maximum induced velocity in the 70% span-wise blade location and the velocity dropped off to zero at the tips. However, the square-tipped rotor appeared to shift the span-wise location of maximum induced velocity further outboard than that of the tapered tip rotor.
Interesting results were also seen when the LL1 duct was added to the rotors.
With the duct in place, it was shown that the span-wise location of maximum velocity
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occurred at the 60% location instead of 70% as was the case for just the isolated rotors. It
was also shown that velocity at the tips in the ducted fan case was not zero. These trends
were attributed to the duct wall impact on the rotor. It was suggested that the rotor tip
vortices are reflected inboard along the blade span after interacting with the duct wall.
The concept of swirl velocity was also introduced. It was seen from the results that a
swirl velocity does exist inside the duct. This had the effect of reducing the velocity below the rotor in the ducted case.
A new procedural method was developed with the 10 series isolated rotors. The velocity was measured just above the rotor as well as just below it. A simple average was then taken to obtain a better estimate of the actual induced velocity. The 10 series experiments also served to explore the usefulness of a new data logging hardware and its software. As a whole these velocity profile experiments were successful in beginning the formulation of an experimental database for ducted fans.
Several important hover thrust experiments were presented in Chapter 5. A detailed description of the facilities used including APB and Hammond was also given.
The procedures used in performing the necessary calibration for each force and moment balance was discussed. The past and present influence coefficient matrices were presented for both balances. The data acquisition system unique to both facilities was also discussed which included instructions on how the data is acquired through the
LabView software. An overview of the post-processing involved during the experiments and non-dimensionalization of aerodynamic data was shown. Finally, the results of each hover thrust experiment were detailed.
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The hover experiments with the LL1 model generated the expected result of
increased thrust with the addition of the duct to the isolated rotor. Overall between two
rotors of varying pitch and tip shape, a thrust augmentation of 30% was achieved when
the duct was included. Similar results were obtained in a separate facility proving that
the experiment was repeatable. This also provided increased experience with alternate
instrumentation.
Finally, the hover experiments with both 10 series models were completed. The
tip clearance study with 10-1 validated the claim that the smaller the tip clearance becomes, the greater the amount of thrust can be produced. The experiment with 10-2 provided an interesting result with the vents opened. Although an overall thrust decrease
of 15% was obtained with the vents opened as opposed to vents closed, the thrust
increased linearly with an increase in rotor tip speed.
The final experimental chapter, Chapter 6, discussed in detail all aspects of the forward flight experiments with the ducted fan models. A description of the design and layout of the wind tunnel facility at APB was given. Previous data obtained to measure the turbulence intensity was presented as a means of validating the quality of the tunnel.
A brief description of the methods used for calibrating the tunnel test section as well as performing calibration checks to the APB balance was also shown. This included several changes to the previous data acquisition system and new LabView computer programs for monitoring the entire experiment. The procedure used to calculate the tare drag was also covered. Finally, before any forward flight data was shown, the non-dimensional aerodynamic coefficients calculated during the experiment were defined. A computer program was also created to post-process all the data obtained from the experiment.
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Initial results compared the measured dimensional lift of an isolated 3 blade rotor to that of each isolated duct, and ducted rotor combination. The dimensional lift was directly compared for each model. Overall, it was seen that the isolated rotor produces substantially more lift than any of the ducts. This was to be expected though, because each single duct is fairly blunt with little to no traditional lifting surfaces. It was distinguished however that model 10-2 with the vents closed produced the greatest amount of body lift than any of the other ducts tested. On average, the ducted fan combination produced almost three times as much lift as the isolated rotor at the same given RPM. When the measured lift of each ducted fan model was non-dimensionalized with wind tunnel speed, it was seen that the lift coefficient was roughly the same for moderate to high wind tunnel speeds. At the lowest wind tunnel speed tested, model 10-2 with the vents open produced the lowest lift coefficient. For all the ducted fans, the lift coefficient (non-dimensionalized with wind tunnel speed) decreased with increasing wind tunnel speed. However, when the measured lift of each ducted fan was non- dimensionalized with rotor tip speed, it was seen that the lift coefficient increased with wind tunnel velocity.
When the drag was measured for each isolated duct, it was shown that 10-2 with the vents open produced the largest dimensional drag force over all wind tunnel speeds.
However, when the rotors were placed in the ducts, it was found that 10-2 with the vents open produced the lowest drag over the majority of all wind tunnel speeds. It was also determined that a portion of the drag measured for the ducted fans was momentum drag since the measured drag increased significantly over that of the isolated duct case. Also, it was seen for each ducted fan at the same RPM that the drag coefficient decreased as the
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forward flight speed increased. When comparing the dimensional drag with lift produced by each ducted fan, it was seen that the aerodynamic efficiency decreased with forward
speed. For the single ducted fan models tested, the drag force dominated the lift force at
wind tunnel speeds greater than 40 ft/sec by almost a factor of two.
The pitching moment about the 3/4 duct chord was also measured for each model.
In general, it was found that a fairly constant nose-up pitching moment was produced for all wind tunnel speeds. Interestingly, it was found that model 10-2 with the vents open produced the lowest pitching moment all wind tunnel speeds. It was actually two times less than that of model 10-1.
Throughout the forward flight experiments, the measured aerodynamic coefficients for each ducted fan model were also presented as functions of advance ratio.
The overall trend was shown that maximum forces and moments, when non- dimensionalized with wind tunnel speed, occurred at very low advance ratios. As the advance ratio increased, these non-dimensional forces and moments appeared to decrease. When the lift was non-dimensionalized with rotor tip speed and presented as a function of advance ratio, the lift coefficient was shown to decrease from its hover value to a minimum at low advance ratios. These trends revealed that potentially the area of most concern in the flight envelope of single ducted lift fan vehicles would be the transition from hover to low-speed forward flight.
Finally, it was concluded that a possible solution to the transition instability was shown through the opening of the vents of model 10-2. Interestingly, the drag and pitching moment were both reduced for advance ratios less than 0.2 when the vents were opened. However, while the drag and pitching moment were reduced, the lift was also
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reduced at these low advance ratios. Thus it appeared that a tradeoff exists such that
some lift must be sacrificed during transition.
7.2 Suggestions for Future Work
Throughout the project, observations were made that could enhance some of the
experiments. For example, some suggestions to the induced velocity experiments include
measuring the velocity above the rotor and past the tip region out to the 110% non-
dimensional radius. Also a measurement of the velocity distribution under the rotor hub
would be beneficial. Both of these extended range measurements would help to further
define the rotor slipstream. It would also be helpful to measure the velocity just beneath
the rotor plane for the ducted cases. This could be accomplished by inserting the mini-
vane anemometer through a hole in the side of the duct.
It was also found that flow visualization techniques could be useful additions to
the wind tunnel experiments. By using a smoke wire or tuft grid, the flow in and around
the duct could be observed. This could be helpful in interpreting the momentum drag
results among others as described in Chapter 6. The use of a smoke wire would
especially be helpful to visualize how the forward vents affect the air flow around the
rotor.
Variations in the forward vent configuration should be performed. It was shown
that fully opening of forward vents of model 10-2 had the potential to reduce drag and pitching moment. It would be interesting to see how partial openings of the vents affect
the overall performance of the ducted fan in forward flight. It is anticipated that the ideal
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vent configuration would reduce both drag and pitching moment, while keeping the loss
in lift as little as possible.
Currently, there are plans to begin a first order study on control vanes placed at
the exit of the 10 series single fan models. As described in Chapter 1, exit control vanes
have the potential to eliminate complicated hub structures that traditional helicopters use
for maneuvering the vehicle. As shown in Figure 7-1, the exit control vanes consist of a
cascade of four symmetric airfoils, each having a 2 inch chord. The vane airfoils are
nominally spaced 1 inch apart and can be positioned at any angle with respect to the exit plane. Experiments will be performed which will attempt to measure any changes in side
force and rolling moment due to exit vane deflection.
Figure 7-1: Exit control vane design for single fan model
Vehicles with twin ducted lift fans have the potential of carrying substantially
more payload than single ducted lift fans. A vehicle, such as the Urban Aeronautics X-
Hawk, with ducted lift fans in the tandem configuration also has the potential to produce
more body lift than that of a single fan vehicle if a proper fuselage is designed. Thus, a
dual ducted fan model was designed and fabricated in a similar fashion to the single fan
models. As shown in Figure 7-2, each duct features the same 9.5 inch diameter as the 10
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series single fan models. The dual fan model essentially consists of two separable parts,
including the isolated rotor frame and the fuselage/ducts. The frame holds the rotors as
shown in Figure 7-3. Notice that the incidence angle, β , of each rotor can be adjusted to
approximately +/- 15 o. The same experiments performed with the single fan models
could be done with both the isolated and ducted dual fan model. Smoke visualization
could also be utilized to help understand the interaction between each fan.
Figure 7-2: Fully assembled dual fan model
Figure 7-3: Isolated dual rotor model and incidence angle adjustment
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It was hypothesized in Chapter 1 that surrounding a rotor with a duct could potentially shield some of the aerodynamic noise produced by the rotor. Preliminary
aeroacoustic experiments were performed with the single fan 10-1 model and 3 bladed
rotor in an anechoic chamber. The radiated noise was measured for both the isolated
rotor and ducted rotor cases, with three microphones on a circular arc of approximately 3
rotor diameters away from the rotor hub. As shown in Figure 7-4, the acoustic spectra
were dominated by tones at the blade passage frequency. It also appears that the ducted
rotor produced a higher level of noise than that of the isolated rotor. It would be
interesting to measure the noise again after installing a sound absorbing lining on the
inside of the duct. More in depth, future aeroacoustic experiments would provide useful
data in addition to the aerodynamic experiments performed in this thesis.
Figure 7-4: Comparison of measured power spectra for the 3-bladed MA ducted and isolated rotors
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Finally, it was observed that the maximum disk loading of the ducted fan models
is quite low compared to other existing ducted fan vehicles as shown in Figure 7-5. The
models produce a disk loading of around 2.0 whereas the disk loading of a traditional
helicopter is approximately 8.0. Thus, a motor upgrade is underway which will
effectively increase the disk loading to compare more favorably to others. Similar three bladed rotors with a higher pitch distribution have also been newly purchased which will
further increase the disk loading.
Figure 7-5: Disk loading comparison of several ducted fans
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Bibliography
[1] Hovey, R. W., “Ducted Fans for Light Aircraft: Analysis, Design, and Construction,” 1982.
[2] Leishman, G. J., “Principles of Helicopter Aerodynamics,” Cambridge University Press, NY, New York, 2006.
[3] Pereira, J. and Chopra, I., “Surface Pressure Measurements on an MAV-Scale Shrouded Rotor in Hover,” American Helicopter Society 62 nd Annual Forum, Phoenix, AZ, May 2006.
[4] Parlett, L. P., “Aerodynamic Characteristics of a Small-Scale Shrouded Propeller at Angles of Attack from 0 o to 90 o,” NACA, Langley Field, VA, 1955.
[5] Martin, P., and Tung, C., “Performance and Flowfield Measurements on a 10-inch Ducted Rotor VTOL UAV,” American Helicopter Society 60th Annual Forum, Baltimore, MD, June 2004.
[6] Anderson, S. B., “Historical Overview of V/STOL Aircraft Technology,” NASA, 1981.
[7] Horn, J. F., “Flight Simulation of Advanced Ducted Fan Air Vehicles,” The Pennsylvania State University, University Park, PA, ONR Presentation, August, 2007.
[8] Sato, T., and Engeda, A., “Design and Performance Analysis of Ducted Fans For Micro UAV Applications,” AHS International Specialists’ Meeting on Unmanned Rotorcraft, 2005.
[9] Abrego, A. I. and Bulaga, R. W., “Performance Study of a Ducted Fan System,” AHS, San Francisco, CA, 2002.
[10] Parlett, L., “Stability and Control Characteristics of a Small-Scale Model of an Aerial Vehicle Supported by Two Ducted Fans,” NASA, Langley Field, VA, 1961.
[11] Yoeli, R., “X-Hawk Wind Tunnel Test #1, Preliminary data and discussion,” September, 2005.
[12] Camci, C., and Akturk, A., “Tip-Leakage Vortex Minimization in Ducted Axial Fans Using Novel Pressure Side Tip Platform Extensions,” Ankarra International Aerospace Conference, Metu, Ankara, September 2007.
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[13] Fleming, J., and Jones, T., Lusardi, J., Gelhausen, P., and Enns, D., “Improved Control of Ducted Fan VTOL UAVs in Crosswind Turbulence,” AHS 4 th Decennial Specialist’s Conference on Aeromechanics, San Francisco, CA, 2004.
[14] Tilford, K., A., “Improvements in the Aerodynamic Testing Capabilities of the PSU Wind Tunnel Facilities,” M.S. Thesis, Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 2007.
[15] Litz, Brian. “Hammond Balance,” Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 2004.
[16] Swan, T. S., “Low Reynolds Number Wind Tunnel Investigation of the PSU 94-097 Winglet Airfoil,” M.S. Thesis, Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 2000.
[17] Germanowski, P. J., “Wind Tunnel Investigation of the Longitudinal Stability of a Ground Effect Vehicle,” M.S. Thesis, Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 2003.
[18] Brophy, C. M., “Turbulence Management and Flow Qualifications of the Pennsylvania State University Low-Turbulence, Low-Speed, Closed-Circuit Wind Tunnel,” M.S. Thesis, Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 1994.
[19] Medina, R., “Validation of the Pennsylvania State University Low-Speed, Low- Turbulence Wind Tunnel Using Measurements of the S805 Airfoil,” M.S. Thesis, Aerospace Engineering Department, The Pennsylvania State University, University Park, PA, 1994.
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Appendix A
A.1 LL1 Series Motor Power Experiment
In order to determine the power requirements of the LL1 14.5” propellers, the power curve for the motor-propeller combination was measured experimentally. To power the motor, a direct current power supply capable of a 50V and 18A range was used. The motor with propeller attached, was held to the edge of a table with a clamp.
The RPM of the propeller was measured by using an infrared light diode as it passed over reflective tape placed on the propeller hub. The frequency between each piece of reflective tape was measured through a spectrum analyzer. Several data points were taken as the voltage was increased to the motor. After plotting power required by the motor against propeller RPM, the power requirement for any RPM could be extrapolated.
A block diagram of the setup is shown in Figure A-1.
Spectrum DC Power Supply Analyzer
Motor Infrared Diode
Power Supply
Figure A-1: Block diagram of power curve experiment with LL1 14.5” rotor with tapered tips
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The same experiment was conducted while varying the gear ratio. The power curve experiment was first run with the 14.5”, tapered tip, propeller directly attached to the motor shaft and hence a gear ratio of 1:1. The propeller RPM could potentially be increased with the use of a gear box. A gear drive was purchased which could vary gear ratios by inserting gear pinions of decreasing size. The smallest gear pinion was used in a power curve experiment which gave a gear ratio of 3.8:1.
A final experiment was performed in which the temperature of the motor was measured for a duration of 20 minutes at maximum propeller RPM. The same experimental setup was used with the addition of an optical temperature gage gun. This experiment was performed with the 14.5”x 11” tapered tip, 4-bladed propeller
Figure A-2 shows the calculated power curve for the 14.5” tapered tip propeller
for a 1:1 gear ratio. As seen from Figure A-2, the power increases with RPM. With a
gear ratio of 1:1, or when the propeller was directly attached to the motor shaft, a
maximum RPM of about 1600 was reached. The maximum RPM was limited by the
current input of the motor. As the voltage was increased, the current reached the cut off
limit of the power supply. It was determined that the propeller puts a high load on the
motor because of its high pitch. A curve fit was also applied to the motor to determine
how much power would be required to operate the propeller at a given RPM. Table A-1
shows the maximum RPM with its corresponding required power and predicted power
required values for extrapolated RPMs.
Several methods were considered for increasing the maximum rotor RPM such as
using a brushed motor with a greater number of turns, implementing a gear drive, or
using a higher RPM motor. It was decided to first try including a gear drive. The gear
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14.5" (11" pitch), Tapered Tip, 4 blade rotor 1:1 Motor Gear Ratio 0.12 *Isolated Rotor
0.1 y = 4E-12x 3.2194 R2 = 0.9938 0.08
0.06
0.04 Power Required [hp] 0.02
0 0 500 1000 1500 2000 Rotor RPM
Figure A-2: Power versus RPM for 14.5” x 11” rotor with a 1:1 motor gear ratio
Table A-1: Extrapolated power required to achieve 10000 RPM with 14.5” rotor and 1:1 gear ratio RPM Power (hp) 1620 0.1 3000 0.63 4000 1.6 5000 3.2 6000 5.8 7000 9.6 8000 14.7 9000 21.5 10000 30.2
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drive purchased allowed the gear ratio to be varied from 2:1, 3.5:1, and 3.8:1. It was
determined that a gear ratio of 3.8:1 would allow for the highest RPM. The measured power curve for the 14.5” tapered tip rotor and motor with a 3.8:1 gear ratio is shown in
Figure A-3. Figure A-3 shows that the maximum RPM achievable was about 2900.
Although this was an increase in RPM from the case when a gear drive was not employed, it still does not fall into the RPM ranges of the experiments conducted by
Parlett[4] and Martin[5] . A curve fit was again applied to the data in order to predict the power required by the motor to operate at certain propeller RPMs.
Table A-2 shows the power required to reach the maximum RPM for the 14.5”, tapered tip, rotor as well as some predicted values of power required for higher propeller
RPM. As seen from the predicted power required values in Table A-2, a fairly large motor would be required in order to operate the 14.5” propeller at 10,000 RPM.
However, 5000 RPM could be reasonably accomplished. Another option of increasing motor RPM was to use a power supply with a larger output current range. Care must be taken however to ensure that the motor does not become too hot to cause motor damage at large values of current. Thus the motor temperature at maximum RPM and hence maximum current was measured for a duration of time.
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14.5" (11" pitch), Tapered Tip, 4 blade rotor 3.8:1 Motor Gear Ratio 0.25 *Isolated Rotor
0.2 y = 4E-11x 2.7713 R2 = 0.9966 0.15
0.1
Power Required [hp] 0.05
0 0 500 1000 1500 2000 2500 3000 3500 Rotor RPM
Figure A-3: Power versus RPM for 14.5” x 11” rotor with a 1:1 motor gear ratio
Table A-2: Extrap olated power required to achieve 10000 RPM with 14.5” rotor and 3.8:1 gear ratio RPM Power (hp) 2940 0.19 4000 0.38 5000 0.71 6000 1.2 7000 1.8 8000 2.6 9000 3.6 10000 4.9
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It was decided to operate the 14.5” rotor with 3.8:1 gear ratio at maximum RPM for 20 minutes. Figure A-4 shows the measured values of temperature. The motor reached a peak temperature of almost 180 oF after the motor was shut off for 1 minute.
The motor then cooled after it was shut down and time increased. Figure A-4 shows that a bigger power supply may not be the answer for increasing propeller RPM. Experiment durations of over 20 minutes at a constant current of about 18 A would damage the motor.
14.5" (11" pitch), Tapered Tip, 4 blade rotor 3.8:1 Motor Gear Ratio 200 *Isolated Rotor 180 160
F] 140 o 120 100 80
Temperature [ 60 40 20 0 0 5 10 15 20 25 30 Time [min]
Figure A-4: Temperature of motor with isolated 14.5” x 11” propeller and a 3.8:1 gear ratio for a duration of 20 minutes
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A.2 Preliminary 10 Series Model Designs
Several designs were considered for the 10 series ducted fan models. Three of these alternate designs are shown as Figures A-5, A-6, and A-7. Each of these designs utilize the two-piece duct construction. The PVC/foam interaction however is different from the chosen design. The alternate designs also feature variations in duct inlet shape.
The models of Figures A-5 and A-6 both feature straight duct exits. The third model shown in Figure A-7 shows the possibility of a duct diffuser that would expand the exit flow. In all the alternate model designs, a foam ring insert was designed that would slip into the inside of the duct. This would allow for variable tip clearances.
Figure A-5: Alternate design of 10 series ducted fan models (baseline)
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Figure A-6: Alternate design of 10 series ducted fan models (change in inlet lip radius)
Figure A-7: Alternate design of 10 series ducted fan models (increased diffuser angle)
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Appendix B
B.1 Setup of Pulleys for Calibration of APB Balance
Figure B-1: Positive pure drag calibration setup
Figure B-2: Positive pure side force calibration setup
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Figure B-3: Negative pure lift calibration setup
Figure B-4: Positive pure roll calibration setup
246
Figure B-5: Positive pure pitch calibration setup
Figure B-6: Positive pure yaw calibration setup
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B.2 APB Balance Calibration Curves
Drag Calibration 2
0 0 2 4 6 8 10 12
2
4
Drag 6 Side Force Output Voltage [V] Lift y = -0.9298x + 0.0102 Roll 8 Pitch Yaw
10 Pure drag loading [lbf]
Figure B-7: APB calibration curves for pure drag loading
Side Force 9 Calibration 8 Drag Side Force 7 Lift 6 Roll Pitch y = 1.0302x + 0.0115 5 Yaw 4 3
Output Voltage [V] 2 1 0 0 2 4 6 8 10 1 Pure side force loading [lbf]
Figure B-8: APB calibration curves for pure side force loading
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(-) Lift Calibration 7 8/26/08 6
5 y = -0.4032x - 0.0461 4 Drag Side Force 3 Lift Roll 2 Pitch Output Voltage [V] Yaw 1
0 18 16 14 12 10 8 6 4 2 0 1 Pure Lift loading [lbf]
Figure B-9: APB calibration curves for pure lift loading
Roll 2 Calibration
0 0 10 20 30 40 50 2
4
6 Drag y = -0.2643x + 0.0108 Side Force Output Voltage [V] 8 Lift Roll 10 Pitch Yaw 12 Pure roll loading [lbf inch]
Figure B-10: APB calibration curves for pure roll loading.
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Pitch 1 Calibration 0 0 20 40 60 80 100 1
2
3 y = -0.0733x + 0.0011 Drag 4 Side Force
Output Voltage [V] Lift 5 Roll 6 Pitch Yaw 7 Pure pitch loading [lbf inch]
Figure B-11: APB calibration curves for pure pitch loading
Yaw 2 Calibration
0 0 10 20 30 40 2
4 y = -0.2705x + 0.0049 6 Drag Side Force Output Voltage [V] 8 Lift Roll 10 Pitch Yaw 12 Pure yaw loading [lbf inch]
Figure B-12: APB calibration curves for pure yaw loading
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B.3 History of Past APB Balance Calibrations
Contained in this section are the influence coefficient matrices, ICMs, from four previous calibrations. These matrices are the slopes of the calibration curves, and not the
inverse of the ICM. Conditions were similar during each calibration. However some of
the calibrations were performed inside and outside of the wind tunnel. This is meant to
give a record of changes in the calibration of the APB balance. The ICMs were all
adjusted for a single gain setting of x10000. The units of the moment coefficients are
V/lb-inch.
Spring 2007
− .0 9300 − .0 0230 − .0 0010 .0 1070 .0 0890 − .0 1280 − .0 0030 .1 0170 .0 0350 .0 2260 − .0 0420 − .0 0130 − .0 0190 − .0 0220 − .0 1850 − .0 0130 − .0 0020 .0 0140 ICM = (B-1) .0 0010 − .0 0020 − .0 0010 .0 2680 .0 0030 .0 0010 .0 0010 − .0 0030 .0 0000 .0 0060 − .0 0750 − .0 0010 .0 0010 .0 0000 − .0 0010 − .0 0020 .0 0030 .0 2500
Fall 2007
− .0 9269 .0 0135 .0 0016 − .0 0002 .0 0002 .0 0004 − .0 0057 − .1 0353 .0 0009 − .0 0009 − .0 0004 − .0 0002 − .0 0007 − .0 0009 − .0 3971 − .0 0006 .0 0003 − .0 0003 ICM = (B-2) .0 1164 .0 1489 − .0 0336 − .0 2696 .0 0055 .0 0049 .0 0305 .0 0481 .0 0023 − .0 0050 − .0 0737 .0 0032 − .0 1473 − .0 0010 .0 0171 − .0 0013 − .0 0010 .0 2469
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Summer 2008
− .0 9298 − .0 0022 .0 0089 .0 0009 .0 0013 .0 0015 − .0 0161 .1 0302 .0 0089 .0 0000 .0 0000 − .0 0051 − .0 0064 − .0 0068 − .0 4032 .0 0011 .0 0006 − .0 0001 ICM = (B-3) .0 1249 − .0 0658 .0 0139 − .0 2643 .0 0043 − .0 0063 .0 0395 − .0 0508 − .0 0025 − .0 0048 − .0 0733 − .0 0043 − .0 0606 − .0 0002 .0 0190 − .0 0005 − .0 0002 − .0 2705
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