In this study, I present a systematic design methodology for a micro fixed-wing drone aimed at low-altitude, low-speed reconnaissance missions. The increasing demand for compact, long-endurance unmanned platforms in both military and civilian sectors has driven the need for efficient fixed-wing drones that operate at very low Reynolds numbers. My work focuses on a flying‑wing layout combined with a twin‑propeller propulsion system, which inherently balances torque effects through counter‑rotating propellers and enhances low‑speed control authority. The entire development follows a parametric approach, starting from mission requirements, through weight estimation, wing geometry determination, and center‐of‐gravity analysis, and concluding with computational fluid dynamics (CFD) simulations to evaluate aerodynamic performance.
The primary mission requirements guiding the design of my fixed-wing drone are:
- Battery‑powered electric motor
- Maximum payload of 45 g
- Endurance greater than 15 min
- Cruise speed of 11 m/s
These constraints define the initial size and weight of the vehicle. For a micro fixed-wing drone, the compactness of the airframe is critical, and a flying‑wing layout is the most suitable because it eliminates the fuselage and tail, reducing overall drag and weight. The twin propellers are placed on the leading edge of the wing to provide a slipstream over almost the entire span, which significantly improves lift and control effectiveness at low speeds. Moreover, the counter‑rotating design cancels the torque reaction, simplifying the flight control system.
Overall Design of the Fixed-Wing Drone
Take‑off Weight Estimation
In the overall design of fixed-wing drones, the take‑off weight is a fundamental parameter that affects all subsequent sizing. The weight breakdown is expressed as:
$$W_0 = W_{\varepsilon} + W_M + W_B + W_{PL}$$
where \(W_0\) is the take‑off weight, \(W_{\varepsilon}\) is the structural weight, \(W_M\) is the power and control system weight, \(W_B\) is the battery weight, and \(W_{PL}\) is the payload weight. For battery‑powered micro fixed-wing drones, statistical data show that the payload fraction is typically around 21% (WPL / W0 ≈ 0.21). Thus:
$$W_0 = \frac{W_{PL}}{0.21} = \frac{45\ \text{g}}{0.21} \approx 213\ \text{g}$$
This value is used as the initial design gross weight.
Wing Geometry Determination
Micro fixed-wing drones operate in the low Reynolds number regime (5×10⁴ to 1.5×10⁵) and usually adopt low aspect ratio (LAR) wings. Based on previous studies by Torres, the aerodynamic characteristics of LAR wings are dominated by aspect ratio, followed by planform shape, and finally Reynolds number. I selected a trapezoidal planform with an aspect ratio of 1.7, a taper ratio of 0.9, and no sweep. The airfoil chosen is E216, which offers a good balance between lift and drag at moderate cruise speeds.
The cruise lift coefficient is estimated as \(C_L \approx 0.37\). Using the lift equation:
$$S = \frac{2 W_0}{\rho V^2 C_L}$$
With air density \(\rho = 1.225\ \text{kg/m}^3\), velocity \(V = 11\ \text{m/s}\), and \(W_0 = 0.213\ \text{kg} \times 9.81 = 2.09\ \text{N}\), the wing area is:
$$S = \frac{2 \times 2.09}{1.225 \times 11^2 \times 0.37} \approx 0.0796\ \text{m}^2$$
The span is then:
$$b = \sqrt{S \cdot AR} = \sqrt{0.0796 \times 1.7} \approx 0.368\ \text{m}$$
From the taper ratio \(\lambda = 0.9\) and area, the root chord \(c_r\) and tip chord \(c_t\) are:
$$c_r = \frac{2S}{b(1+\lambda)} = \frac{2 \times 0.0796}{0.368 \times (1+0.9)} \approx 0.246\ \text{m}$$
$$c_t = \lambda c_r \approx 0.221\ \text{m}$$
The mean aerodynamic chord (MAC) is:
$$\bar{c} = \frac{2}{3}c_r\frac{1+\lambda+\lambda^2}{1+\lambda} \approx 0.234\ \text{m}$$
Stall speed is a critical parameter for launch and recovery. For my fixed-wing drone, the stall speed was estimated from the maximum lift coefficient (assumed \(C_{L,max} \approx 1.0\)):
$$V_S = \sqrt{\frac{2W_0}{\rho S C_{L,max}}} \approx 5\ \text{m/s}$$
This low stall speed allows safe hand‑launching.
Center of Gravity Estimation
I determined the center of gravity (CG) location by summing the moments of all components. The table below lists the masses and their CG positions relative to a reference point at the nose.
| Component | Mass (g) | X‑CG (mm) | Y‑CG (mm) | Moment X (N·m) | Moment Y (N·m) |
|---|---|---|---|---|---|
| Wing | 43.16 | 141.7 | 16.4 | 0.0599 | 0.0069 |
| Vertical tail | 2.63 | 146.6 | 12.4 | 0.0038 | 0.0003 |
| Fuselage structure | 4.21 | 82.4 | 6.65 | 0.0034 | 0.0003 |
| Payload | 45 | 27 | 10 | 0.0132 | 0.0049 |
| Motor (×2) | 11.2 | 45 | 4 | 0.0049 | 0.0004 |
| ESC | 24 | 60 | 2 | 0.0141 | 0.0005 |
| Battery | 33 | 40 | −10 | 0.0129 | −0.0032 |
| Flight controller | 5.58 | 200 | 10 | 0.0109 | 0.0005 |
| Servos | 18 | 118 | 9 | 0.0208 | 0.0016 |
| Ballast / spare | 26.22 | – | – | – | – |
| Total | 213 | 0.1441 | 0.0123 |
The CG location along the X‑axis is:
$$X_{CG} = \frac{0.1441\ \text{N·m}}{0.213\ \text{kg} \times 9.81\ \text{m/s}^2} \approx 0.0748\ \text{m} = 74.8\ \text{mm}$$
Relative to the mean aerodynamic chord (MAC = 234 mm), the CG is at 23.8% MAC. This position ensures static longitudinal stability for the flying‑wing configuration.
Summary of Wing Parameters
| Parameter | Value |
|---|---|
| Wing area (S) | 0.0796 m² |
| Span (b) | 0.368 m |
| Aspect ratio (AR) | 1.7 |
| Taper ratio (λ) | 0.9 |
| Sweep angle | 0° |
| Mean aerodynamic chord (MAC) | 0.234 m |
| Design lift coefficient (CL) | 0.37 |
| Stall speed (VS) | 5 m/s |
Aerodynamic Analysis Using CFD
After completing the conceptual design, I built a full three‑dimensional model of the fixed-wing drone in CATIA. The model includes the flying wing, twin vertical stabilizers placed near the wingtips (which also act as winglets to reduce induced drag), and streamlined nacelles for the motors. A half‑model was used in the CFD simulation due to symmetry, and a hybrid mesh with refined cells around the leading edge, trailing edge, and tip regions was employed. The simulations were performed with the Spalart‑Allmaras turbulence model, which is suitable for attached and mildly separated flows at low Reynolds numbers.
Figure above shows the CAD model of the fixed-wing drone used in the CFD analysis. The compact flying‑wing layout with twin propellers is clearly visible.
Force and Moment Results
The aerodynamic coefficients were computed for angles of attack ranging from 0° to 26°. The lift coefficient (\(C_L\)), drag coefficient (\(C_D\)), and lift‑to‑drag ratio (\(L/D\)) are summarized in the following table and graphs.
| Angle of Attack (°) | \(C_L\) | \(C_D\) | \(L/D\) |
|---|---|---|---|
| 0 | 0.12 | 0.032 | 3.75 |
| 2 | 0.24 | 0.055 | 4.36 |
| 4 | 0.35 | 0.071 | 4.93 |
| 6 | 0.46 | 0.089 | 5.17 |
| 8 | 0.56 | 0.108 | 5.19 |
| 10 | 0.65 | 0.128 | 5.08 |
| 12 | 0.74 | 0.150 | 4.93 |
| 14 | 0.83 | 0.175 | 4.74 |
| 16 | 0.91 | 0.203 | 4.48 |
| 18 | 0.98 | 0.235 | 4.17 |
| 20 | 1.05 | 0.271 | 3.87 |
| 22 | 1.10 | 0.310 | 3.55 |
| 24 | 1.06 | 0.350 | 3.03 |
| 26 | 0.98 | 0.392 | 2.50 |
The lift curve shows a nearly linear increase up to about 20°, with a maximum \(C_L\) of approximately 1.10 at 22°, after which stall occurs. This stall behavior is typical for low‑aspect‑ratio fixed-wing drones, where vortex lift and flow separation are delayed. The maximum lift‑to‑drag ratio is about 5.2 near 8° angle of attack. Although this value is modest compared to high‑aspect‑ratio aircraft, it is acceptable for the low Reynolds number regime where viscous effects dominate.
Pressure Distribution
Contours of pressure coefficient on the upper surface reveal a strong suction peak near the leading edge at moderate angles of attack. At high angles (above 22°), a large separation bubble appears on the upper surface, leading to lift loss. The tip vortices are weak because of the low aspect ratio, but the vertical stabilizers help to partially recover some of the vortex energy. The pressure distribution also confirms that the twin‑propeller slipstream, although not simulated in this steady‑state analysis, would further energize the boundary layer and delay separation, a phenomenon that will be studied in future work.
Conclusions
I have successfully completed the conceptual design and aerodynamic analysis of a micro fixed‑wing drone tailored for low‑speed, long‑endurance missions. The key accomplishments and findings are:
- The flying‑wing layout with a low aspect ratio of 1.7 and zero sweep provides a compact, lightweight platform. The twin counter‑rotating propellers effectively cancel torque and improve low‑speed lift and control, addressing a core challenge in micro fixed‑wing drone design.
- The take‑off weight is estimated at 213 g, and the wing area is 0.0796 m², yielding a wing loading that allows a stall speed of 5 m/s. The CG is located at 23.8% MAC, ensuring static longitudinal stability.
- CFD simulations show that the fixed-wing drone exhibits a maximum lift coefficient of 1.10 at 22° angle of attack and a peak lift‑to‑drag ratio of 5.2 near 8°. These results confirm that the design is aerodynamically viable in the low Reynolds number range.
- The pressure contours indicate that flow separation becomes significant above 20°, which will be mitigated in later design iterations by optimizing the wing twist or adding leading‑edge devices.
This study provides a solid foundation for further development of micro fixed‑wing drones. Future work will include time‑accurate CFD with rotating propellers to capture slipstream effects, wind‑tunnel validation, and flight testing of a prototype. The methodology presented here can be readily adapted to other mission requirements, demonstrating the flexibility of the parametric design approach for fixed-wing drones.