Development of an in-flight thrust measurement system for ... - AIAA ARC

Development of an in-flight thrust measurement system for UAVs Andrew Gong∗, Hugh Maunder† and Dries Verstraete‡ The University of Sydney, Sydney, NSW, 2006, Australia

Small electric unmanned aerial vehicles have traditionally lacked the capability to directly measure the thrust of the propulsion system. Without this thrust measurement capability, drag estimation of the aircraft through flight tests such as glide tests becomes highly inaccurate due to the additional drag of the wind-milling propeller and the high sensitivity of small vehicles to gusts. Direct measurement of thrust enables an accurate determination of the aircraft drag polar from flight tests, as well as the performance and efficiency of the propulsion system itself. This paper presents the design and development of a prototype lightweight, accurate in-flight thrust measurement system for use in small unmanned aerial vehicles. Calibration results are given along with data from wind tunnel and flight tests. Estimates of aerodynamic performance parameters are then calculated using measurements from the thrust measurement system, showing the ability to use this system for more accurate airframe performance evaluation of small unmanned aircraft.

Nomenclature CAD IFTMS RC UAV

Computer aided design In flight thrust measurement system Remote control Unmanned aerial vehicle

I.

Introduction

Small unmanned aerial vehicles (UAVs) are finding increasing use in a range of industries throughout the world, from agriculture through to commercial mapping as well as Defence operations.1–6 With these increased uses comes the desire for improved performance and range to maximise the effectiveness and utility during a particular UAV mission. Arguably the most important factor in the performance of an aircraft is the drag polar. The drag polar and thrust form the basis for which all performance characteristics of the aircraft may be calculated.7 Traditional performance evaluation methods require extensive wind tunnel testing with a model at a range of airspeeds and angles of attack.8 This is a time-consuming process and usually ignores the significant effects of propeller slipstream on lift and drag. Similarly, glide tests face inaccuracies due to the additional drag of the wind-milling propeller.7 Each of these methods has limitations in the accuracy of drag estimation and aircraft performance evalution. An in-flight thrust measurement system (IFTMS) would enable the direct measurement of thrust during all flight conditions. This would provide a simple and accurate solution to measuring both aircraft and propulsion system performance.9 This paper presents the design, construction and testing of a prototype in-flight thrust measurement system in a small electric UAV. The design and test of the strain gauge mount is outlined, followed by ∗ Doctoral Candidate, School of Aerospace, Mechanical and Mechatronic Engineering, The University of Sydney, NSW, Australia, 2006, AIAA Student Member † Undergraduate Student, School of Aerospace, Mechanical and Mechatronic Engineering, The University of Sydney, NSW, Australia, 2006, AIAA Student Member ‡ Senior Lecturer, School of Aerospace, Mechanical and Mechatronic Engineering, The University of Sydney, NSW, Australia, 2006, AIAA Senior Member

1 of 10 American Institute of Aeronautics and Astronautics

calibration results and integration of the thrust measurement system into a UAV airframe. Finally, wind tunnel and flight test results are presented with an estimation of aircraft performance parameters based on the in-flight thrust measurements.

II.

Design of the Strain Gauge Mount

Two types of thrust cell mounting devices were developed and tested in the prototyping stage. The two types of thrust cell mounting devices considered were the hinge type and linear bearing type. A description of each of the types along with advantages and disadvantages is given below. A.

Hinge Prototype

A hinge type mounting system was tested based on the concept of David Hares’ flying thrust cell.10 Two firewalls are used in this design, with one firewall used as the mounting point for the motor and the second firewall mounted directly to the airframe. A hinge was used to join the bottom edge of the two firewalls, and the strain gauge mounted along the upper edge to measure the moment about the hinge, and hence the thrust generated by the propulsion system. A top-view and side-view of the prototype are presented in Figure 1.

(a) Top-View

(b) Side-View

Figure 1: Hinge Bearing Mounting Prototype

B.

Linear Bearing Prototype

A linear bearing design was also prototyped due to concerns about induced moments about the hinges of the hinge type design. The extraneous moments can be caused by turbulence experienced in flight, pitch rate accelerations or yaw rates resulting in propeller gyroscopic moments. A linear bearing does not suffer from issues of induced moments about the hinge. However, there are other shortcomings related to friction in the bearing and the impact on the linearity and accuracy of the load cell measurements. A prototype linear bearing was constructed based on four linear bearings between the aircraft firewall and motor mount. This set-up allows fore and aft movement in the aircraft longitudinal axis, but restricts rotational movement about this axis. The prototype set-up is shown in Figure 2.

III.

Strain Gauge Calibration

To verify the thrust measurement system, the strain gauge output of both prototypes was measured and compared with the load cell output of an RCBenchmark 1850 Dynamometer. A series of test weights were loaded onto each prototype and the output voltage reading recording. These results comparing the in-flight 2 of 10 American Institute of Aeronautics and Astronautics

(a) Top-View

(b) Side-View

Figure 2: Linear Bearing Mounting Prototype

strain gauge output and the existing RCBenchmark load cell are shown in Figure 3.The strain gauge design using the hinge showed excellent linearity over a range of test loads. In comparison, the linear bearing showed poor linearity at test loads below 0.3 kg. Thus, the hinge type system demonstrated better performance over the range of test loads compared to the linear bearing system.

Figure 3: Strain Gauge Calibration Based on these test results, the hinge type system was selected as it was established to be more accurate over a wide test range. Although there is the possibility of extraneous moments, these can be minimised through the test regime by taking measurements during steady cruise conditions. Furthermore, the hinge brings the advantages of mechanical simplicity, reduced weight and scalability.

IV.

Final Thrust Measurement Design

With the selection of the hinge type thrust measurement system, a final design was made incorporating modifications to improve the overall accuracy. A high quality spring loaded hinge was selected to minimise hinge friction and reduce mechanical clearances over the prototype design. This design also reduced the impact of other frictional losses such as flexing of the motor wires. The CAD design is illustrated in Figure 4a, with the final thrust cell installation shown in Figure 4b. The final thrust measurement cell was tested and calibrated using a series of static weights. Ten full static loading cases were performed using calibrated weights up to a maximum of 2 kg. The results of this

3 of 10 American Institute of Aeronautics and Astronautics

(a) CAD Drawing of Final Thrust Cell

(b) Thrust Sensor Installed in Airframe

Figure 4: Airframe and Thrust Sensor Installation

calibration can be seen in Figure 5. There is excellent linearity and repeatability over the full range of measured forces.

Figure 5: Final Static Load Calibration

V.

Airframe Integration

The in-flight thrust measurement system was installed on a twin-engine RC model Cessna 310 by Dynam RC Products displayed in Figure 6. This model has a wingspan of 1.28 m with a take-off weight of 1.75 kg. The thrust measurement system is designed to be mounted directly to the existing mounting bracket located on the aircraft firewall. The engine cowl required some minor lengthening to ensure sufficient clearance with the thrust measurement system in place. Figure 4b shows the thrust sensor installed on one motor of the aircraft and mounted to the airframe. Specifications for the Cessna 310 model can be found in Table 1. The aircraft was provided with a stock propulsion system consisting of 2 x Detrum 1100 kV motors and 2 8x6” three-bladed propellers. This motor/propeller combination was found to produce approximately 7.3 N of static thrust at full power. A series of wind tunnel and flight tests were performed using the thrust measurement system to characterise the propulsion system performance.

VI.

Wind Tunnel Results

Wind tunnel testing was undertaken in the 7 x 5 ft low speed wind tunnel located at the University of Sydney. This closed return wind tunnel has an airspeed operating range of 2-42 m/s. The Cessna 310 model was mounted onto an ATI Industrial Automation Mini45 Force/Torque Sensor.

4 of 10 American Institute of Aeronautics and Astronautics

Figure 6: Cessna 310 model in 7’x5’ wind tunnel

Table 1: Model Cessna 310 Specifications Specification Wingspan Fuselage Length Mass Aspect Ratio Average Chord Wind Loading Wing Area

Value 1.28 m 1.1 m 1.2 kg 7.9 0.1616 m 58 g/dm2 0.2069 m2

This 6 axis load cell was used to measure aerodynamic forces and moments experienced in the wind tunnel. It was also used as a check on the measurements recorded by the thrust measurement system mounted to the model aircraft. Due to the lack of room in the model and the desire to use the same aircraft for both wind tunnel and flight testing, the load cell was mounted to an external plate on the underside of the UAV. This affects the accuracy of the results and causes an increase in the installation drag. The aircraft drag polar was determined for the Cessna 310 model at a series of different recorded tunnel speeds, using the method outlined by Traub.8 There is good agreement of the drag polar between the different airspeeds determined in the wind tunnel for a range of airspeeds from 6 to 20 m/s. In flight thrust measurement system validation Wind tunnel testing of the model Cessna 310 equipped with the thrust measurement system provided the opportunity to validate the performance of the system. The drag polar was used to predict the drag at a range of airspeeds at zero angle of attack. In conjunction, the propulsion system of the wind tunnel model was set so that the longitudinal force on the load cell read zero. This corresponds to the thrust provided by the propulsion system exactly equalling the drag of the aircraft. The results of these were then plotted simultaneously on Figure 8. As shown, there is clear agreement between the wind tunnel load cell and the aircraft mounted thrust measurement system. This validated the performance of the thrust measurement system in atmospheric conditions close to those expected in flight.

VII.

Flight Test Results

A series of flight tests were performed to validate the in-flight thrust measurement system. These flight test were performed at a range of throttle settings and with different flight conditions including steady-level cruise, level acceleration, steady climb, steady descent and glide test. An example of one flight test is shown in Figures 9 and 10. Figure 9 displays the measured thrust and RPM for each of the motors and propellers.

5 of 10 American Institute of Aeronautics and Astronautics

Figure 7: Cessna 310 drag polar

Figure 8: Load cell vs thrust measurement system

As expected, the measurements of both motors show a close correlation.

Figure 9: Propeller Thrust and RPM Based on the thrust and RPM measurements of the motor and propeller, the propulsive power generated can also be determined. The propulsive power out is given by P = T V where T is the thrust and V is the airspeed. The input power delivered by the battery and the output propulsive power are graphed in Figure 10.

6 of 10 American Institute of Aeronautics and Astronautics

Figure 10: Power Provided and Propulsive Power Output

Three different methods were used during the flight testing to collect aircraft performance data. Each of the test flights was segmented into sections containing the specific flight manoeuvre required for each method. These methods are presented in following subsections. A.

Level Acceleration

The drag polar of the aircraft can be represented by the zero-lift drag coefficient (CD0 ) and the Oswald span efficiency e by the quadratic drag polar equation: CD = CD0 + K(CL − CL0 )2

(1)

1 where K = πARe . The performance characteristics of CD0 and e are then used to determine the propulsive power requirement,

Preq =

1 3 2W 2 ρV SCD0 + K 2 ρV S

(2)

where V is airspeed, S the wing area, W the weight and ρ air density. Rearranging this equation by multiplying by V produces an equation with V 4 as the independent variable. Preq · V =

1 4 2W 2 ρV SCD0 + K =a·V4+f 2 ρS

(3)

CD0 can then be determined by the gradient of the graph CD0 =

2·a ρV 2 S

(4)

where a is the linear fit gradient. The Oswald efficiency e can be determined from the y-intercept f using e=

2W 2 ρSπARf

(5)

McCrink used level acceleration runs for this method as they have a range of excess power and velocity values.11 Figure 11 shows the results from five level acceleration runs. The curve fit yielded the results of CD0 of 0.0551 and e of 0.6527 with a y-axis standard deviation of 384. There is significant scatter visible in Figure 11, and this is due to the manner in which the flight test was conducted. The aircraft was manually piloted making it difficult to achieve truly straight level flight. Furthermore, there was significant wind during the period of flight testing. B.

Power-Off Glides

Power-off glides were used for evaluating aircraft performance using the method outlined by Lowry.12 The glide angle was determined using γ = sin−1 ( Vδhδt ), where a GPS unit was used to determine the change in altitude. The zero-lift drag coefficient can then be calculated using CD0 =

W sinγ ρV 2 S

7 of 10 American Institute of Aeronautics and Astronautics

(6)

Figure 11: Straight-and-Level Acceleration Run

and the Oswald efficiency determined by e=

4CD0 πAR · tan2 γ

(7)

In this flight regime, there is significant drag due to the windmilling propellers. Corrections were included in the calculations using the drag values measured by the thrust measurement system to more accurately estimate the airframe drag. The mean of two glide tests yielded values of 0.0545 for CD0 and 0.236 for e. The correct and uncorrected zero-lift drag coefficients are presented in Figure 12. This glide test was performed at an airspeed of 14 m/s.

Figure 12: Glide Test Results

C.

Straight and Level Tests

During steady level flight conditions, the equilibrium conditions of lift equalling weight and thrust equalling drag enable the direct calculation of performance parameters from the linear drag polar. A series of straight and level flight sections was analysed from the flight testing, and the mean result used to generate the linear curve fit (Figure 13). The fit was found to yield a Oswald efficiency of e = 0.687 with an intercept value of CD0 = 0.0542 and standard deviation of 0.0107. A comparison of each of the performance estimation methods results is given in Table 2. The calculated value of CD0 for the wind tunnel test is significantly higher than the flight test methods. This is due to the exterior mounting arrangement of the load cell, which results in significantly higher drag. The three flight test methods are in close agreement for the value CD0 , with a calculated range of 0.0542-0.0551. This represents a maximum difference of 1.6% between the three methods. For the Oswald span efficiency e, the level acceleration and straight-and-level tets are in close agreement, with values within 6% of the other. The calculated value of e for the glide test differs significantly. This is 8 of 10 American Institute of Aeronautics and Astronautics

Figure 13: Steady Level Test Results

likely due to inaccuracies in the altitude estimation using the GPS sensor. The use of a differential pressure sensor for more precise altitude estimation would likely improve the accuracy of this method. Table 2: Calculated Parameters Parameter CD0 e

Wind Tunnel 0.0805 0.307

Level Acceleration 0.0551 0.653

Straight and Level 0.0542 0.687

Glide 0.0545 0.236

During the steady-level flight conditions, the coefficient of drag can be determined by equating lift to weight and thrust to drag. The thrust is directly measured by the in-flight thrust measurement system and averaged for each section operated at a constant airspeed. Figure 14 shows the calculated drag polar, while Figure 15 shows the calculated drag coefficient as a function of the airspeed. A quadratic fit for the data was determined using Matlab. Error bars are also shown on Figure 14 representing the standard deviation in CD . The estimated minimum drag lift coefficient CLmd is 0.33, which correlates well with the value of 0.35 estimated in the wind tunnel test. The curve fit in Figure 15 suggests a CDmin of 0.0505 at approximately 18 m/s.

Figure 14: Flight Test Drag Polar

9 of 10 American Institute of Aeronautics and Astronautics

Figure 15: Drag Coefficient vs Airspeed

VIII.

Conclusion

This paper has introduced the motivation behind the development of an in-flight thrust measurement system for small electric UAVs. An overview of the design and test of the prototype system is presented, along with calibration and airframe integration results. Wind tunnel and flight test results are presented, and analysis to determine aircraft aerodynamic parameters is also outlined. The paper has shown the benefits of an in-flight thrust measurement system in performance estimation of small UAVs. Future work includes the integration of an autopilot for more precise flight manoeuvres and an improved hinge mount design to minimise extraneous load cases from pitch manoeuvres and gyroscopic effects.

References 1 Gundlach, J., Designing unmanned aircraft systems: a comprehensive approach, AIAA Education Series, American Institute of Aeronautics and Astronautics, Reston, VA, USA, 1st ed., 2012. 2 Gong, A., Palmer, J., Brian, G., Harvey, J., and Verstraete, D., “Performance of a hybrid, fuel-cell-based power system during simulated small unmanned aircraft missions,” International Journal of Hydrogen Energy, Vol. 41, No. 26, 2016, pp. 11418–11426. 3 Gundlach, J. and Foch, R. J., Unmanned Aircraft Systems Innovation at the Naval Research Laboratory, Library of Flight, American Institute of Aeronautics and Astronautics, Reston , VA, 2014. 4 Gong, A. and Verstraete, D., “Development of a dynamic propulsion model for electric UAVs,” APISAT 2015 - 7th Asia-Pacific International Symposium on Aerospace Technology, 2015, pp. 206–217. 5 Verstraete, D., Cazzato, L., and Romeo, G., “Preliminary design of a fuel-cell-based hybrid-electrical UAV,” 28th Congress of the International Council of the Aeronautical Sciences 2012, ICAS 2012 , Vol. 1, 2012, pp. 422–431. 6 Verstraete, D., Gong, A., Lu, D. D. C., and Palmer, J. L., “Experimental investigation of the role of the battery in the AeroStack hybrid, fuel-cell-based propulsion system for small unmanned aircraft systems,” International Journal of Hydrogen Energy, Vol. 40, No. 3, 2015, pp. 1598–1606. 7 Kimberlin, R. D., Flight Testing of Fixed-Wing Aircraft, AIAA Education Series, American Institute of Aeronautics and Astronautics, Reston, VA, USA, 1st ed., 2003. 8 Traub, L. W., “Validation of endurance estimates for battery powered UAVs,” Aeronautical Journal, Vol. 117, No. 1197, 2013, pp. 1155–1166. 9 Bronz, M. and Hattenberger, G., “Aerodynamic Characterization of an Off-the-Shelf Aircraft via Flight Test and Numerical Simulation,” AIAA Flight Test Conference, 2016. 10 Hares, D., “Building a Thrust Test Fixture,” Report, 2011. 11 McCrink, M. H. and Gregory, J. W., “Flight Test Protocol for Electric Powered Small Unmanned Aerial Systems,” AIAA Atmospheric Flight Mechanics Conference, 2014. 12 Lowry, J. T., Performance of Light Aircraft, American Institute of Aeronautics and Astronautics, Inc.,, Reston, Va, USA, 1999.

10 of 10 American Institute of Aeronautics and Astronautics