See discussions, stats, and author profiles for this publication at: https://www.researchgate.net/publication/340852614 Propulsion selection method using motor thrust table for optimum ο¬ight in multirotor aircraft Conference Paper in AIP Conference Proceedings · April 2020 DOI: 10.1063/5.0004809 CITATIONS READS 0 1,281 6 authors, including: Harish Mahatma Putra Muhamad Rausyan Fikri Miota Internasional Teknologi Sampoerna University 5 PUBLICATIONS 4 CITATIONS 28 PUBLICATIONS 48 CITATIONS SEE PROFILE R. Dimas Pristovani Electronics Engineering Polytechnic Institute of Surabaya 16 PUBLICATIONS 57 CITATIONS SEE PROFILE Some of the authors of this publication are also working on these related projects: Hybrid VTOL Unmanned Aerial Vehicle View project Vehicle Dynamical Systems View project All content following this page was uploaded by Harish Mahatma Putra on 06 May 2020. The user has requested enhancement of the downloaded file. SEE PROFILE Propulsion selection method using motor thrust table for optimum flight in multirotor aircraft Cite as: AIP Conference Proceedings 2226, 060008 (2020); https://doi.org/10.1063/5.0004809 Published Online: 22 April 2020 H. M. Putra, M. R. Fikri, D. P. Riananda, G. Nugraha, M. L. Baidhowi, and R. A. Syah ARTICLES YOU MAY BE INTERESTED IN Smart pulley workflow in delivery drone for goods transportation AIP Conference Proceedings 2226, 060010 (2020); https://doi.org/10.1063/5.0006800 AIP Conference Proceedings 2226, 060008 (2020); https://doi.org/10.1063/5.0004809 © 2020 Author(s). 2226, 060008 Propulsion Selection Method Using Motor Thrust Table for Optimum Flight in Multirotor Aircraft H. M. Putra1, a), M. R. Fikri2, b), D. P. Riananda1, c), G. Nugraha1, d), M. L. Baidhowi1, e), and R. A. Syah3, f) 1 2 Drone Research Squad, IoT Research Department, PT. Bukalapak.com, DKI Jakarta, Indonesia. Head of IoT & Physics Lab, Information System, Sampoerna University, DKI Jakarta, Indonesia 3 Head of IoT Research Department, PT. Bukalapak.com, DKI Jakarta, Indonesia. a) Corresponding author: [email protected] b) [email protected] c) [email protected], [email protected] d) [email protected] e) [email protected] f) [email protected] Abstract. One of the problems when designing an aircraft is flight time estimation. When we design the aircraft, in general, calculating the maximum takeoff weight becomes the priority. The longer flight time required by the aircraft, the more batteries should be added. More batteries mean the payload capacity of the aircraft is decreased. On the other hand, adding more weight will increase the requirement of using more powerful motors. We need to take into our account, that using powerful motors should be calculated carefully because the current consumption is also increased. This problem is circling in the same situation, without any real solution to overcome the problem. As unmanned aircraft has become a critical part of human life for example in surveillance, agricultural, and medicine delivery mission, obtaining the optimum and efficient flight time is vital. In this research, we propose a novel approach of maximum takeoff weight calculation that is needed when designing the aircraft by looking at the complete map of motor specifications. The goal of this paper is to help any aircraft user from hobbyist to professional to determine whether their aircraft design is feasible. INTRODUCTION Nowadays, Unmanned Aerial Vehicles (UAVs) are common technology in the world. There are two main categories based on its component where a lift is produced which is fixed-wing and rotary-wing.1 The fixed-wing aircraft produces its lift by non-movable wing attached to its body. Meanwhile, the rotary-wing produces its lift from moving wing in a circular motion known as a propeller. The rotary-wing has many advantages such as vertical takeoff and landing, low speed and stationary flight, and no need for runway.2 There are subcategories for rotary-wing based on its number of rotors: single-rotor and multirotor UAV. Multirotor UAV has many rotors, for example, four, six, and eight. Multirotor UAV uses its thrust to lift and to control its attitude. However, multirotor has one disadvantage, it requires high power consumption. When designing the multirotor, generally start from defining maximum take-off weight (MTOW) and payload. If we are not considering the weight of the propulsion system and batteries earlier, there will be a later problem. One of the examples is when the actual flight time does not meet the specification, adding one or more batteries to make it fly longer is not a solution. The additional battery will add overall weight and reducing payload capacity. Meanwhile, changing motor with a new powerful motor will increase current consumption and obviously, larger battery capacity is needed. This process of selecting the payload and battery will occur repeatedly and become the problem that should have a solution since it is time-consuming and costly. 7th International Seminar on Aerospace Science and Technology – ISAST 2019 AIP Conf. Proc. 2226, 060008-1–060008-7; https://doi.org/10.1063/5.0004809 Published by AIP Publishing. 978-0-7354-1985-8/$30.00 060008-1 In order to tackle this repeated problem, in this study, we propose a method to measure and estimate the capability of the motors then produce information about the recommendation of payload capacity on the Thrust Table. We elucidate the detail information start from calculating the motor thrust data in section 2, the steps of creating the Thrust Table in section 3, then the discussion of the Thrust table in section 4 with the conclusion and future work in section 5. MOTOR THRUST DATA In this method, first, we collect every motor thrust and current consumption data. We are using thrust and current consumption with specific motor types and propeller combination from T-Motor as our research data. Those data also can be acquired by using a thrust benchmark system.3 FIGURE 1. Example curve. Thrust and current relation of T-Motor MN501s kv3601 Thrust and Current Relation The thrust data and current consumption can be fit into a curve. For every different combination, e.g. the same motor but for different propeller or vice versa, need a new curve. From that curve, we can make an approximation using second-order polynomial. The general form of the polynomial equation can be expressed4 as Eq. (1). π(π₯) = ππ₯ 2 + ππ₯ + π (1) Where π(π₯) is the value of Y-axis, the output of the polynomial function π, π, π are constants for second order polynomial π₯ is value of X-axis In this case, π(π₯) is current in Ampere and π₯ is thrust in kg. Because the curve intersects with point (0,0) so there will be no c. So, we can write the relation between thrust and current as Eq. 2. This approach makes a possibility to estimate motor thrust and current relation without knowing the actual motor scientific model. πΌ = ππ₯ 2 + ππ₯; 0 ≤ π₯ ≤ π₯πππ₯ Where πΌ is current consumption in Ampere (A) π, π are constants from second order regression π₯ is thrust in kilogram (kg) π₯πππ₯ is maximum thrust (kg) 1 Data source: http://store-en.tmotor.com/goods.php?id=697 (17 July 2019) 060008-2 (2) Where π and π are constants, π₯ is the amount thrust needed, and πΌ is current consumption. There is a limit for thrust, it is a maximum value from specification data. We should make a limit in later formula. Calculating Required Battery From the known current usage and flight time required, we can use Eq. (3) to estimate the required battery capacity. The 1 Ampere-hour means that the battery will supply 1 Ampere for an hour.5 Multiplying total current consumption and time will result in battery capacity and total current consumption can be acquired by multiplying current usage per motor with the number of motors. π‘ πΆπππ = 1.25 β π β πΌ β πππ (3) 60 Where πΆπππ is battery capacity in Ampere-hour (Ah) π is number of motors in multirotor πΌ is current usage per motor in Ampere (A) π‘πππ is flight time required in minute The current I from the Eq. (2), can be inserted into Eq. (3) in Ref. 6. The result is the required battery capacity in Ampere-hour (Ah). The unusable and spare battery capacity also needs to be considered. When the capacity below 10%, the battery should stop discharging.7 For safety reason and possible emergency case, 20% of total battery capacity should be reserved. So, extra value 125% of battery capacity needs to be expected. One thing to note, this calculation does not include the battery discharge characteristic.8 Most batteries have less capacity when discharged with a higher current. Lithium batteries are well known to have a quite linear characteristic between capacity and discharge rate. So, this method uses an assumption of the linearity of lithium battery. After we get the required battery capacity, we need to calculate the number of parallel cells. There is an input value of real battery capacity. Simply using Eq. (4). ππππ = ππππππππ ( πΆπππ πΆππππ ) (4) Where ππππ is number of parallel cells πΆπππ is required battery capacity (Ah) πΆππππ is real battery capacity (Ah) The total battery cells required can be calculated in Eq. (5) using result from Eq. (4). ππ‘ππ‘ = ππππ β ππ ππ (5) Where ππ‘ππ‘ is total number of cells ππππ is number of parallel cells ππ ππ is number of series cells The value of ππ ππ is available from motor specification. From Eq. (4) the weight of the battery can be estimated by multiplying a total number of cells and individual battery cell weight. Where π€πππ‘ is total estimate battery weight (kg) ππ‘ππ‘ is total number of cells π€ππππ is individual cell weight (kg) π€πππ‘ = ππ‘ππ‘ β π€ππππ 060008-3 (6) Calculating Propulsion Weight The propulsion weight includes a battery, motor, and propeller. This will simply add the weight of motor, ESC, and propeller and multiply by number or rotor. Then add the weight of the battery. ππ = π€πππ‘ + π(π€πππ‘ + π€ππ π + π€ππππ ) (7) Where ππ is the propulsion weight π€πππ‘ is the total weight of battery π€πππ‘ is the total weight of individual motor π€ππ π is the weight of individual ESC π€ππππ is the total weight of individual propeller BUILDING THE TABLE This table requires iterating calculation. First, we specify an array of take-off weight value. For every value, calculate it by inserting take-off weight value into Eq. (2). Then insert motor current consumption from Eq. (2) into Eq. (3), until we have the total weight of the propulsion system and battery ππ from Eq. (7). πππ = πππ − ππ (8) Where πππ is the available weight for airframe and payload πππ is the total takeoff weight ππ is the total weight propulsion system With Eq. (9), we can calculate the remaining weight for airframe and payloads, including avionics. We need to iterate this calculation for the next takeoff weight value, then do the same calculation for all available motors. This calculation will be simpler when we use a spreadsheet. Now, after we have the whole map of possible propulsion configuration, the last step is we need to make sure that the combination has valid value and enough thrust to weight. Thrust to weight ratio is the comparison between the maximum thrust of all propulsions and weight of the aircraft.9 π πππ = ππππ₯ πππ ππππ₯ = π β π‘πππ₯ (9) (10) Where π πππ is thrust to weight ratio ππππ₯ is total maximum thrust πππ is takeoff weight π is number of rotors π‘πππ₯ is maximum thrust per rotor, from motor specification Some calculations can have heavier propulsion system than the takeoff weight, resulting in negative airframe weight. Some calculations also can result in heavier takeoff weight than the maximum thrust of the propulsion. Those calculations need to be excluded and marked as impossible. 060008-4 Zone Overpower Safe Critical Impossible TABLE 1. Zoning of the thrust table. Meaning Color π > π πππ₯ Yellow π πππ < π ≤ π πππ₯ Green 1 < π ≤ π πππ Red π ≤ 1 or ππ > πππ Grey Therefore, we need to consider thrust to weight ratio. So, there will be 5 zones in the table as mentioned in table 1. The first zone is where the thrust to weight ratio R less than 1 or the weight of the propulsion system ππ is heavier than specified take-off weight πππ , which means the design is impossible. When the total weight of the aircraft exceeds the maximum force of the propulsion system, the propulsion system is unable to lift the aircraft. The other meaning of the impossible zone is when maximum take-off weight is below the weight of the propulsion system. That means we need a negative weight to make design possible. FIGURE 2. The thrust table in colored spreadsheet The second zone is the safe zone where the design is possible. In that zone, the thrust to weight ratio R is between specified thrust to weight minimum π πππ and thrust to weight maximum π πππ₯ . Typically, thrust to weight ratio11 is 2, but it depends on the required maneuverability. The other zones are the overpower zone where the thrust is considered too much and the critical zone where the thrust is doubtfully able to lift the aircraft. In the critical zone, the propulsion system theoretically can lift the aircraft but doesn’t have sufficient power for maneuvering. All those ratio limits are depending on the required maneuverability. In that case, those values may vary in each design. 060008-5 RESULT AND DISCUSSION Several factors that have not been taken into consideration yet, such as price, propeller size, battery configuration, and availability. Some motors could be not in stock or use a high voltage battery that requires high voltage ESCs. Based on table 1, there are several options can satisfy our design requirement, marked with green cells in specific takeoff weight column. The numbers on the cells are indicating the available weight for airframe and payload. We can simply choose the largest number in the column if we do not consider another factor such as propeller size and cost. The green area is the safe zone, as described in table 1, all the possible combinations that matched our requirement. The yellow area marked thrust-to-weight ratio is in overpower zone. The red area is the critical zone and the grey area marked the impossible zone because of negative available weight or thrust to weight ratio is less than 1. All the green zone in a column are our options. The next step we have to do is designing the airframe with a requirement of the weight less than a number in a green cell including its payload. If the design of airframe is not satisfying the calculation, we need to look at another takeoff weight column or another green cell. Study Case In our case, we need to design a multirotor capable to lift 5kg payload and a minimum of 20 minutes of flight time for a delivery operation. The first thing we do is input 20 minutes of flight time, then choose our battery. For the battery, we are using with lithium-ion NCR18650GA because of high density and readily available on the market. Each NCR18650GA cell has 3.5Ah capacity and 48gram weight. We have no constraint for propeller size. We look for green cells with a value of at least 5 kg. We discover that hexacopter is a better option for our case than a quadcopter. FIGURE 3. Options for our design From the result in Fig. 4, we find out that T-Motor MN501s kv300 with 22” propeller is the best option. When we design with 13kg maximum takeoff weight (MTOW), it can lift 8.53kg weight excluding propulsion and battery. With calculation from Eq. (4), we need 6 series and 7 parallels of NCR18650GA. By removing 5kg from 8.53kg, we have 3.53kg for airframe and additional components. From that result, we know that the safe limit of the chosen configuration is 18kg. The more MTOW, the more batteries need to be added for the same flight time. If we choose 14kg MTOW, we have 10.5kg for payload and airframe but we need 8 parallels of NCR18650GA. If we design with MTOW above 18kg, the aircraft may not lift properly or difficult to maneuver once airborne. CONCLUDING REMARKS AND FUTURE WORKS From this experiment, we can conclude that our method is able to help make decisions easier by measuring motor capability based on the table. This method can provide several options for the most optimum options from the available propulsion system. For further work and improvement, the method can be combined with a thrust benchmark system to collect actual motor data. Thrust benchmark system is a tool to obtain the propulsion model10 but this method only needs thrust and current information from test bench. Another thing to note, this calculation currently uses the assumption that multirotor is in hovering mode in the same air pressure as the test bench and does not include current consumption when maneuvering. ACKNOWLEDGEMENT The author would like to acknowledge that this work was supported by PT Bukalapak.com. 060008-6 REFERENCES 1. H. González-Jorge, J. Martinez-Sánchez, M. Bueno, and P. Arias, 2017 “Unmanned aerial systems for civil applications: A review,” in MDPI Drones 1 (MDPI, 2017), pp. 1-19 2. S. Bouabdallah, P. Murrieri, and R. Siegwart, 2004 “Design and control of an indoor quadcopter” IEEE International Conference Robotics and Automation (ICRA) (IEEE, 2004), pp. 4393-4398 3. G. Hattenberger, A. Drouin, and M. 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