In the development of high-altitude long-endurance fixed-wing UAVs, the power system plays a pivotal role in ensuring mission success. The matching between the airframe, engine, and propeller is critical for achieving optimal performance across various flight phases. This paper presents a systematic methodology for designing and evaluating the power system of a fixed-wing UAV, starting from mission profile design, through engine and propeller selection, to final matching analysis. Key parameters such as thrust requirements, power demand, propeller diameter constraints, and efficiency curves are considered. The results demonstrate that the selected power system meets all mission demands, providing a reliable framework for fixed-wing UAV power system design.
Introduction
Unmanned aerial vehicles (UAVs) have seen rapid development in recent years, with their power systems being a focal point of research. For fixed-wing UAVs with a takeoff weight under 1000 kg and a service ceiling below 8000 m, piston engines are widely adopted due to their simplicity and lightweight characteristics. Among piston engines, two-stroke aviation piston engines are particularly favored for their structural simplicity and high power-to-weight ratio. The propeller, driven by the engine, generates the thrust necessary for flight. Therefore, the proper matching of the fixed-wing UAV, its engine, and its propeller is essential for fulfilling mission requirements.
Previous studies have focused primarily on engine-propeller matching, often neglecting the interaction with the airframe. This work addresses that gap by proposing a holistic matching method. We begin by defining the mission profile of a fixed-wing UAV, followed by calculations of thrust and power demands. Then, an engine is selected based on peak power requirements, and a propeller is chosen considering tip speed limits and engine characteristics. Finally, a detailed matching analysis verifies that the selected components work harmoniously under all operational conditions.
Mission Profile Design for a Fixed-Wing UAV
The mission profile of a fixed-wing UAV graphically represents the flight trajectory and tasks. For a typical reconnaissance fixed-wing UAV, we design a vertical flight profile consisting of takeoff, climb, cruise, descent, and landing phases. The fixed-wing UAV operates under standard atmospheric conditions, with a normal takeoff weight, covering a range of 200 km at altitudes between 3 km and 7 km. Table 1 summarizes the mission requirements for different flight states.
| Flight State | Horizontal Speed (m/s) | Vertical Speed (m/s) | Transmission Efficiency (%) | Power Demand for Generator (kW) |
|---|---|---|---|---|
| Idle | – | – | – | – |
| Engine Warm-up | – | – | – | 1 |
| Takeoff at Sea Level | 0–30 | 12 | 98 | 4 |
| Cruise at 3 km | 44 | 0 | 98 | 3 |
| Maximum Speed at 3 km | 54 | 0 | 98 | 3 |
| Landing | 40 | -7 | 98 | 2 |
| Cruise at 5 km | 44 | 0 | 98 | 3 |
| Maximum Flight Altitude (7 km) | 44 | 0 | 98 | 3 |
The mission profile includes climb and cruise phases where the fixed-wing UAV must generate sufficient thrust to overcome drag and weight components. During climb, the required thrust is given by:
$$
T_{\text{req}} = D + W \sin\alpha
$$
where \(T_{\text{req}}\) is the required thrust, \(D\) is the drag, \(W\) is the aircraft weight, and \(\alpha\) is the climb angle. The lift generated by the wings is:
$$
L = W \cos\alpha = C_L \frac{1}{2} \rho V^2 S
$$
and the drag is:
$$
D = C_D \frac{1}{2} \rho V^2 S
$$
Combining these, the required thrust becomes:
$$
T_{\text{req}} = W \left( \frac{C_D}{C_L} \cos\alpha + \sin\alpha \right)
$$
For level flight (cruise), \(\alpha = 0\), so:
$$
T_{\text{req}} = \frac{W}{K}
$$
where \(K = C_L / C_D\) is the lift-to-drag ratio. The power required from the engine is then:
$$
P_e = \frac{T_{\text{req}} V}{\eta} = \frac{W V}{\eta K}
$$
where \(\eta\) is the propeller efficiency, initially assumed to be 0.7 for preliminary selection. Using these formulas, we compute the thrust and power demands for each flight state.
Engine Selection for the Fixed-Wing UAV
For fixed-wing UAVs flying below 300 km/h, a piston engine combined with a propeller is the most economical choice. The engine must provide sufficient power at altitude. The power output at altitude can be estimated using the empirical formula:
$$
N_H = N_0 \left[ 1.11 \frac{P_H}{P_0} \sqrt{\frac{T_0}{T_H}} – 0.11 \right]
$$
where \(N\) is power, \(P\) is atmospheric pressure, \(T\) is temperature, subscript \(H\) denotes altitude condition, and subscript \(0\) denotes sea level condition.
Based on the mission requirements, the maximum power demand occurs during maximum speed flight at 3 km. Using the fixed-wing UAV structural parameters (weight, wing area, drag polar), we compute the required engine power. The preliminary calculation gives a required maximum power of 57.5 kW. Considering generator loads and transmission losses (Table 1), the engine must provide at least 62 kW. Therefore, we select a spark-ignition aviation piston engine with a maximum power of 78 kW at 6000 rpm, which provides a safety margin. Table 2 lists the key specifications of the selected engine.
| Parameter | Value |
|---|---|
| Maximum Power | 78 kW |
| Rated Speed | 6000 rpm |
| Reduction Gear Ratio | 2.43:1 |
| Type | Two-stroke, spark-ignition |
Propeller Selection for the Fixed-Wing UAV
The propeller must be carefully chosen to match both the engine and the fixed-wing UAV. Two critical constraints govern propeller diameter: tip speed limitations and power absorption. The tip speed \(V_{\text{tip}}\) is:
$$
V_{\text{tip}} = \sqrt{(\pi n D_p)^2 + V^2}
$$
The Mach number at the tip \(M_{\text{tip}} = V_{\text{tip}} / c\) must remain below 0.75 to avoid shock waves, where \(c\) is the speed of sound. Additionally, from noise considerations, tip speed should not exceed 213 m/s. Using the engine’s maximum propeller speed (after reduction: 6000 / 2.43 ≈ 2469 rpm), and the maximum flight speed of 54 m/s at 3 km (where speed of sound is about 328 m/s), the maximum allowable propeller diameter is calculated to be 1.99 m.
Another constraint comes from empirical relationships between power and diameter. For a three-blade propeller:
$$
D_p = 0.5 \sqrt[4]{P_e}
$$
With \(P_e = 78\) kW, the suggested diameter is about 1.55 m. Therefore, the actual diameter should not exceed this value. We survey available propellers and select three candidates with diameters of 1.52 m, 1.45 m, and 1.40 m, labeled Propeller 1, 2, and 3 respectively. All are three-blade fixed-pitch propellers.
Matching Analysis of the Fixed-Wing UAV Power System
The matching analysis involves three steps: engine-propeller matching, propeller-fixed-wing UAV matching, and overall verification. We use the engine’s external characteristic curve and the propeller’s power absorption curve. For a fixed-pitch propeller, the absorbed power is proportional to the cube of rotational speed:
$$
P_p = A \cdot n^3
$$
The ideal operating point occurs where the engine power equals the propeller absorbed power. Figure 4 in the original study (conceptually described here) shows the engine curve and the three propeller curves. The maximum power demand for the fixed-wing UAV is 62 kW, chosen as the match point. Propeller 1 and Propeller 2 both intersect the engine curve, while Propeller 3 does not, meaning it cannot absorb enough power. The intersection for Propeller 1 is closer to the 62 kW point, indicating the best match. Therefore, Propeller 1 (diameter 1.52 m) is selected.
Next, we verify that the selected propeller provides sufficient thrust for all fixed-wing UAV flight states. The propeller performance is governed by the advance ratio:
$$
J = \frac{V}{n D_p}
$$
Using the propeller’s characteristic curves (similar to Figure 5 and 6 in the original study), we obtain the power coefficient \(C_P\) and thrust coefficient \(C_T\) as functions of \(J\). The propeller thrust is:
$$
T_p = C_T \rho n^2 D_p^4
$$
and the propeller efficiency:
$$
\eta_p = \frac{T_p V}{P_p}
$$
For the cruise condition at 3 km (V=44 m/s, engine speed 5400 rpm, propeller speed 2222 rpm), we compute \(J = 0.81\), \(C_P = 0.22\), and from the curves, \(\eta_p = 0.61\) and \(C_T = 0.248\). The resulting thrust is 780.5 N, while the fixed-wing UAV requires only 578.3 N for cruise (calculated from \(T_{\text{req}} = W/K\)). Thus, the selected propeller meets the requirement. Table 3 summarizes the thrust calculations for all critical flight states.
| Flight State | Required Thrust (N) | Available Thrust (N) | Status |
|---|---|---|---|
| Takeoff at Sea Level | 1120 | 1250 | OK |
| Cruise at 3 km | 578 | 781 | OK |
| Maximum Speed at 3 km | 720 | 890 | OK |
| Climb at 3 km | 950 | 1050 | OK |
| Cruise at 5 km | 510 | 690 | OK |
| Cruise at 7 km | 440 | 600 | OK |
The matching analysis confirms that the fixed-wing UAV power system, consisting of the selected 78 kW engine and the 1.52 m three-blade propeller, operates efficiently across the entire mission envelope. The engine operates near its optimum power region during cruise, and the propeller absorbs the available power without exceeding tip-speed limits.
Conclusion
This paper presents a comprehensive methodology for designing and matching the power system of a fixed-wing UAV. Starting from mission profile design, we derived thrust and power requirements for each flight phase. An engine was selected based on peak power demand, and a propeller was chosen by balancing tip-speed constraints and power absorption. The final matching analysis verified that the power system meets all demands, including takeoff, climb, cruise, and maximum speed conditions. The approach emphasizes the importance of considering the fixed-wing UAV, engine, and propeller as an integrated system. The results provide a practical reference for engineers developing fixed-wing UAV propulsion systems, ensuring reliable and efficient performance throughout the mission.