1. Introduction
Advances in electronics, materials, and production techniques have allowed for the manufacture of powerful and compact electric motors, high power electrochemical batteries, integrated sensor arrays, and versatile flight controllers suitable for multirotor unmanned aerial vehicle (UAV) applications. However, the key downside of multirotor UAVs is their inherently limited flight endurance and useful range which are directly related to the aircraft take-off mass (aircraft and payload mass combined) and battery charge and energy capacity. A fully electric multirotor powered by batteries, currently being the most widespread multirotor design, is typically characterized by a flight autonomy between 15 and 30 min, with a flight duration of 90 min currently being achieved by high-performance battery-powered multirotor UAVs.
The initial motivation of this research was to alleviate the aforementioned drawbacks of energy sources based purely on electrochemical batteries, by utilizing an alternative propulsion system combining an internal combustion engine (ICE) coupled with the electricity generator (EG) and augmented with a lithium-polymer (LiPo) battery, thus forming a hybrid power unit suitable for multirotor UAV use. Due to the specific energy density of gasoline fuel (about 12 kWh/kg) being practically two orders of magnitude greater than the specific energy density currently available with state-of-the-art lithium polymer (LiPo) batteries (ranging about 0.2–0.5 kWh/kg) [
1], it is expected that the use of a hybrid power unit can facilitate considerable improvements in flight endurance and useful range of such multirotor UAV.
Hybrid propulsion is nowadays typically found in road vehicles and rail transport, as shown by analyses presented in [
2,
3] Since land-based vehicle motion is two dimensional, with gravity-related constraints being much less emphasized compared to those of aerial vehicles, hybrid propulsion system implementation for multirotor-based UAVs would pose a significantly greater challenge, which represents an additional motivation to investigate the proposed research topic.
Hence, the problem of multirotor UAV flight autonomy is actively investigated field and different power system configurations are currently being researched with the aim of satisfying low power consumption, low mass, and high output power density requirements. In a majority of cases, hybrid propulsion system research is conducted for standard fixed-wing and vertical takeoff and landing (VTOL) fixed-wing hybrid [
4,
5], and some other types such as dirigible UAV [
6,
7]. A detailed analysis of synergetic effects for various classes of UAVs is given in [
8] by applying the multi-objective optimization, where results have shown that hybrid-electric configuration has the potential to give a strong contribution to aircraft performance. Moreover, by investigating the behavior of different power sources within the UAV hybrid-electric propulsion system through simulations in order to predict the hybrid electric power system behavior using bench tests, the feasibility and efficiency of the onboard UAV power system can be assessed before the final flight test phase [
9]. Also, a comparison of five different UAV power-train options has been investigated in [
9] using simulations. These alternatives included (i) a free-piston engine with integrated linear EG, (ii) advanced lithium-ion batteries, (iii) ICE with the embedded rotary EG, (iv) a parallel hybrid power-train configuration with ICE, and (v) the proton exchange membrane fuel cell.
Hybrid-electric propulsion system using an ICE has been researched in [
10], with reference [
11] further analyzing some realistic challenges related to ICE use in hybrid-electric propulsion systems. These included acoustic noise and associated mechanical vibrations, engine cooling issues, and their implications to the operation of a compact power unit comprising a small–scale ICE equipped with suitably sized EG. In particular, it has been shown that for such a small ICE, a powerful and complex vibration pattern can be obtained, and it cannot be easily related to engine crankshaft rotation or linear piston motion.
Reference [
12] investigates the potential of hydrogen fuel cell stacks as an alternative power source to ICE in small UAVs. The investigation was based on the commercial “Aeropack” hybrid power supply consisting of a fuel cell stack and a battery pack. The functionality of such a fuel-cell battery hybrid power system has been successfully demonstrated during a flight test of the target prototype UAV.
References [
13,
14] show that photovoltaic power is mostly suitable for fixed-wing UAVs, primarily as an auxiliary power source.
Additionally, UAV hybrid power source research and development efforts have resulted in several patents [
15,
16,
17], which, unfortunately, do not present detailed information about control system design, which is crucial for the implementation of such hybrid power sources. Some other literature sources, such as reference [
18], present the development of a 1 kW hybrid electric power train along with its characterization and performance testing. An interesting approach is given in reference [
19], dealing with a dual power design concept, where UAV motion control is performed through commonly used brushless motors driving a set of auxiliary propellers, while the major portion of UAV lifting power is provided by a gasoline engine-powered main propeller system.
Taking into account the above issues related to specific energy density, mass, ease of operation and performance, and price and availability, hybrid propulsion of multirotor UAV’s based on an ICE coupled to an EG and utilizing a battery energy storage system as auxiliary power supply should provide many attractive research challenges in the field, while simultaneously opening new research frontiers which might ultimately improve knowledge levels in multiple research sub-areas [
20].
Even though internal combustion engines are frequently used for light model aircraft propulsion, their controls are not a frequently discussed topic in the more recent research literature. In order to model the highly nonlinear nature of the internal combustion engine dynamics, the so-called mean-value engine model (MVEM) is typically used as a basis for engine speed controller design, wherein PI and PID controllers can be optimized for different engine speed-torque operating points [
21].
For example, in [
22] a PI controller with feed-forward load compensator is considered for engine speed control within the UAV hybrid propulsion, which may utilize available generator-based measurements (current and voltage) to establish a speed-sensorless feedback loop, or, alternatively, the readily available low-precision Hall-sensor position measurements can be used for that purpose [
23], with Kalman filtering used in both cases to provide a relatively smooth and precise engine speed estimate.
In reference [
24] the three different hybrid-electric power-plant configurations are considered and dynamic models are derived, with flight dynamics performance testing carried out on a UAV prototype.
A good review of current hybrid-electric propulsion systems (HEPS) for fixed-wing aircraft can be found in [
25]. A generic hybrid propulsion system based on DC-AC inverter plus PMSM machine has been presented and modeled in reference [
26], wherein a BLDC generator plus active front end rectifier has been considered as a viable alternative to electric power source hybrid-propulsion UAV in [
27]. In the latter case, using PI and PID controllers has shown favorable engine speed and DC bus voltage control system performance.
An alternative to active rectification of speed-dependent generator voltage has been researched in [
28], wherein a parallelized array of low-cost DC/DC power converters and battery energy storage have been used in conjunction with passive diode rectifier to maintain the fixed voltage at the common DC bus within the UAV.
In order to determine the optimal hybrid power-train configuration and required hybrid propulsion system components, a parameter matching method has been proposed in [
29] wherein the correlation between rotor-based propulsion power demands and the hybrid power system requirements have been identified, with test results confirming the accuracy of the proposed parameter matching method.
Having the aforementioned in mind, this paper proposes a straightforward approach to modeling, identification, and control system design for multirotor UAV power unit system based on ICE coupled with the EG and augmented with the LiPo battery energy storage system. To validate the proposed approach, the custom-built hybrid power unit prototype with dedicated control unit has been built and used to conduct both the process model identification and control system verification.
The paper is organized as follows:
Section 2 gives an overview of the proposed hybrid propulsion system topology, along with the corresponding mathematical models of individual subsystems, in particular, the battery energy storage system, the ICE, and the EG set with a full-wave rectifier.
Section 3 presents the control system design for the ICE-EG set based on the so-called damping optimum criterion, along with key simulation results.
Section 4 presents the results of comprehensive experimental validation of the proposed hybrid propulsion system design. Discussion of key findings presented in the paper and the comparative assessment of hybrid vs. purely electrical propulsion system benefits is presented in
Section 5, whereas concluding remarks are presented in
Section 6.
4. Experimental Verification of DC Bus Voltage Feedback Control System
This section presents the design and development of the proposed hybrid propulsion system experimental setup (schematically shown in
Figure 14a) which was subsequently tested under realistic DC bus electrical load conditions to verify its functionality and validate the simulation model.
4.1. Experimental Hybrid Power Unit Realization
The experimental setup (a photograph is shown in
Figure 14b) consists of the frame that holds all components together, the ICE-EG set that is connected to the common axle using a claw coupling and equipped with mechanical dampeners and springs, a LiPo battery, graduated cylinder tank with the capacity of 0.5 L, and various electronic circuitry including the two microcontroller boards. Separate microcontrollers are used for the ICE throttle control and data acquisition/telemetry tasks in order to avoid possible data acquisition and code execution bottlenecks. Both microcontrollers are programmed and monitored through a host portable computer running MATLAB/Simulink™ software environment.
Table 3 lists the key parameters of individual components used in the setup, along with brief descriptions of these components.
The microcontroller running the PID control algorithm is equipped with a DC bus voltage sensor connected to the appropriate analog-to-digital converter input, this providing the DC voltage measurement needed to establish the feedback loop for the PID controller (as elaborated in the previous chapter). Moreover, this microcontroller also provides the actuator reference for the ICE throttle valve actuator (in the form of a suitable PWM signal). Thus, when DC bus voltage excursion due to electrical load change is detected through the change of feedback signal, the PID controller adjusts the ICE throttle PWM reference, consequently correcting the engine rotational speed and torque, i.e., the overall power output of the engine-generator set feeding the common DC bus.
Any load excess that cannot be compensated for by the DC bus voltage PID con-troller is taken over by the battery (see
Figure 1), whose discharging is mandated by the perceptible DC bus voltage drop sufficient to forward bias the blocking diode. In this way, straightforward passive energy management is implemented, which is desirable from the standpoint of overall control system robustness and redundancy.
The second microcontroller used solely for data acquisition only is equipped with current sensors for the EG current, battery current, DC bus total current, rpm hall sensor, and DC bus voltage sensor. Data acquisition is executed within the MATLAB/Simulink™ software environment, using the so-called simulation model “external mode” execution thus facilitating real-time telemetry.
In order to have a safe and reliable connection between the ICE and EG, a claw coupling of sufficient torque rating is used. It is secured using an appropriate adapter with a conical hole on the engine side to connect it to the engine shaft (
Figure 15a). On the EG side, the coupling must be fixed using a screw characterized by sufficient strength to withstand the load torque variations due to the stroke-based operation of the ICE (see locking pins in
Figure 15b).
Figure 16 represents the validation of the drive simulation model shown in
Section 2.3. Different modes of hybrid drive operation were considered, from the idle throttle, acceleration, and deceleration of the engine. It is shown that the proposed simulation model captures all of the dominant engine-generator system modes quite well, and as such can be used for the synthesis of the DC bus voltage control system.
4.2. Engine Fuel Consumption
The fuel consumption was measured in engine idle regime first, followed by five characteristic operating points that corresponding to EG power output range between 300 to 1700 W. EG unit power was subsequently dissipated by the power resistor network, which was also simultaneously measured by means of a suitable DC wattmeter (see
Figure 17a).
Each test was conducted under steady-state load conditions, by keeping the engine in the particular operating regime for over 5 min. For each steady-state load case, initial and final volumes of fuel were measured using the graduated cylinder tank (as shown in
Figure 17a). Each test was repeated five times to obtain a reliable fuel consumption estimate. The final results of averaged fuel consumption for each operating regime are shown in
Figure 17b. The presented results indicate that fuel consumption characteristic is practically linear with respect to the EG unit electrical load.
4.3. Hybrid Power Unit Measurements
In order to test the functionality of the proposed hybrid power system concept under realistic load conditions, a system was tested in several operating regimes corresponding to no-load (idling), low load, medium load, and high load operating modes. In all experiments, the voltage reference is set to 50 V. The DC bus voltage PID controller was implemented in the C programming language complying with the proposed PID algorithm structures presented in
Section 3.
A stepwise load is chosen for the testing of the hybrid power system transient and steady-state performance for the following reasons: (i) it provides the most abrupt load profile, usable for stringent testing of the control system transient performance, including testing for possible saturation effects and stability issues (closed-loop damping issues), and (ii) it can be easily related to a sudden vertical ascending maneuver, because in that case all propeller drives are suddenly loaded with an almost equal constant (stepwise) load which is maintained until the desired flight level is reached.
Each test was repeated five times and consists of the following DC bus loads emulated by the power dissipation resistor network equivalent resistance:
4 Ω load (low load);
2 Ω load (medium load);
1.33 Ω load (high load);
1 Ω load (peak load).
During tests, ICE is initially held in idle conditions for approximately 5 min (engine warm-up period), followed by the DC bus load being stepped up and down, with each load step lasting several seconds to record the corresponding load-on and load-off DC bus system voltage and current transients. Offsets in current measurements are due to non–ideal sensor characteristics, notably emphasized while measuring low (near–zero) currents, which would be removed by periodic recalibration of current sensors in actual applications. Since the tests conducted herein were rather short, these offsets were simply subtracted offline after the measurements were carried out.
Low load responses are presented in
Figure 18a,b. Initially, during the engine run-up from idle speed to the chosen operating point around 10,000 rpm, the battery provides the DC bus load (139–143 s). As the engine speed (rpm) reaches around 10,000 rpm, the EG takes over the load, and the battery current drops to zero. DC bus voltage remains stable near the target value of 50 V throughout the experiment, without significant voltage drop excursions.
Results for medium-high and high load regimes are shown in
Figure 18c,d. Engine ramps up its rpm to approximately between 13,500 rpm and 14,000 rpm. As was the case for low load, the battery initially supplies DC bus load until the EG takes over, with the average generator load being around 30–35 A, thus corresponding to 1500–1750 W of power consumption at the load dissipation resistor network. Again, DC bus voltage is maintained near the target value of 50 V without significant voltage drop excursions.
Results for the case of peak load are shown in
Figure 18e,f. Due to limited ICE-EG set power output, the battery needs to supply the additional current load (approx. 16 A) to the DC bus when the EG current limit of 30 A is reached. In this case, the overall hybrid power system is characterized by maximum power production, amounting to 2400–2500 W of total power output. The DC bus voltage again remains very stable near the reference voltage, without significant voltage excursions under abrupt load current changes.
5. Discussion
Based on the obtained results, it can be concluded that a functional hybrid power system has been developed comprising an ICE-EG set and LiPo battery, with the output DC voltage being fully supported by the power output of the ICE-EG set up to DC bus electrical loads amounting to 1700 W. For larger loads magnitudes, the battery pack is capable of supplying the additional peak load, so a total of 2200 W (up to 2.5 kW in best cases) can be obtained from the proposed hybrid power system. The EG is capable of continuously delivering up to 1200 W without significant heating, with increased heating of the EG (up to 85 °C) being noticed when the power demand exceeds 1500 W, primarily due to EG being fully enclosed, which may prevent adequate cooling at higher loads.
The hybrid power supply overall efficiency can be estimated in the following manner:
where
is the mechanical power at the ICE output, and
is the electrical power transferred from the generator to the load (i.e., power at the dissipation resistor grid). Since the engine can produce 1.8 Nm of torque at 10,500 rpm, what would correspond to the mechanical power of 1980 W. Under these conditions, electrical power dissipated that the load is 1500 W, and the overall efficiency of the hybrid power system is estimated to the aforementioned value of 0.75.
The hybrid drive vs. conventional drive efficiency analysis can be determined based on the energy obtained over a particular reference time interval. In order to equate the energy capacity of the battery and the hybrid drive energy delivery capacity in a straightforward manner, the equivalent hybrid drive energy production in Watt-hours was calculated based on the recorded data and compared with values of the energy capacity of typical batteries for UAV applications. In particular, the internal combustion engine fuel consumption is about 30 mL/min under sustained power production of 1500 W at the generator (corresponding to medium to high generator loads), thus resulting in hourly fuel consumption of about 1.8 L of fuel. This corresponds to 1500 Wh of available energy at the generator, with the overall mass of the hybrid power unit (engine, generator, rectifier, fuel) amounting to 6 kg.
On the other hand, the UAV battery pack under examination is characterized by 400 Wh of energy capacity (see results of battery identification section), with the total mass of the battery and the supporting equipment of about 3 kg. Thus, in order to match the energy output of the hybrid power unit, a total of 4 battery packs are needed, resulting in an overall battery system mass of 12 kg. It is, therefore, easily concluded that the ICE-EG hybrid power unit has a gravimetric energy density of 300 Wh/kg, while the energy-equivalent battery pack has an energy density of about 140 Wh/kg. This clearly shows that the energy (i.e., flight) autonomy of a UAV equipped with a hybrid power unit is significantly improved in comparison to the battery-only ”benchmark” case (see
Figure 19).
It is important to point out that when designing the experimental setup, the authors have mounted the ICE-generator set on rubber dampeners (so-called rubber bobbins or bushings). An estimate of the vibrations magnitude was done by using vibration tests that were implemented in the UAV multirotor “Pixhawk 2” controller unit and “Mission Planner” software. It turns out that the vibration modes that occur during the operation of the ICE-based hybrid propulsion are outside the frequencies that would affect the sensors (inertial unit) and, thus, create navigation problems and stability problems for the aircraft in general. This fact inspires confidence that it should not be too complicated to decouple the hybrid drive from the frame of the aircraft and thereby reduce the impact of vibrations generated by ICE.
6. Conclusions
Based on the results of computer-based simulations and experiments conducted using the hybrid power unit test bench, and the comparative assessment of the conventional battery power unit with the proposed hybrid power unit, several clear benefits of hybrid power supply are identified, which are given below:
Hybridization of the conventional ICE–EG set results in a stable power source is obtained which is characterized by the gravimetric energy density which is two times higher compared to a purely battery-based power supply;
For the aforementioned increase in the gravimetric energy density using a hybridized ICE + EG power unit, the overall mass of the hybrid power system is two times smaller when compared to the comparable battery-based system.
According to the obtained results that indicate the maximum stable power obtained from the hybrid propulsion, in the future the plan is to build a multirotor aircraft (presumably a quadcopter) with a takeoff mass of 10–12 kg with highly efficient 22–24 inch agricultural-UAV propellers, with typical hovering regime power requirements of approximately 1000–1200 W, which would be provided for by a hybrid drive, and an additional power margin of 300–500 W, wherein any power demand above 1500 W would be covered by the battery.
In the future, it would be interesting to explore hybrid power supplies featuring energy recovery, where the battery would be recharged if the generated power of the hybrid drive allows charging of the battery, by means of simultaneous DC bus voltage and battery current control, which may require additional power electronic systems in the form of bidirectional DC/DC power converter.
Of particular interest would be the application of an active rectifier, where a stable DC bus voltage could be generated, independent of the ICE speed, which motivates further research of fuel consumption-optimal ICE control.