Building a Homemade Subsonic Wind Tunnel: High School Aerodynamics Project

The Reynolds number measures the ratio between inertial forces and viscous forces in fluid flow. For meaningful aerodynamic comparisons between scale models and full-scale vehicles, similar Reynolds numbers must be maintained. This requires adjusting wind speed proportionally to achieve equivalent flow characteristics, as simple linear scaling of dimensions does not produce equivalent aerodynamic behavior.

Similarity variables are non-dimensional numbers that must remain constant when scaling aerodynamic problems to ensure accurate comparison between different scales. The Reynolds number (Re = ρLV/μ) compares inertial forces to viscous forces, indicating how important viscosity is to a specific flow problem. Preserving Reynolds number during scaling ensures consistent pressure distribution, viscous drag characteristics, flow separation behavior, and laminar/turbulent flow regimes between scaled and full-scale configurations. Proper non-dimensionalization allows coefficients to remain constant regardless of object size or velocity.

For dynamic similarity between prototype (air) and model (water), Reynolds numbers must match: Re_p = Re_m → (ρ_p V_p D_p)/μ_p = (ρ_m V_m D_m)/μ_m. Given prototype: V_p = 100 km/h, D_p = 1 m, ρ_p = 1.22 kg/m³, μ_p = 1.78 × 10⁻⁵ Pa·s; model: D_m = D_p/4 = 0.25 m, ρ_m = 1000 kg/m³, μ_m = 1 × 10⁻³ Pa·s. Solving for V_m gives approximately 3.85 m/s. The force ratio is (ρ_m/ρ_p)(V_m/V_p)²(A_m/A_p) = (1000/1.22)(3.85/100)²(1/16) ≈ 4.

The Reynolds number is a dimensionless quantity defined as Re = (ρ × V × L) / μ, where ρ is fluid density, V is velocity, L is a characteristic length (such as wing chord), and μ is dynamic viscosity; it enables engineers to scale down aircraft models while maintaining identical airflow behavior by matching the Reynolds number through adjustments to velocity, density, or viscosity, allowing accurate wind tunnel testing regardless of whether the model is tested in air or water.

The Reynolds number (Re = ρVD/μ) is a dimensionless number that characterizes the flow regime. The instructor explains that for dynamic similarity between a model and prototype, the Reynolds numbers must be equal. This ensures that the flow patterns around the model and prototype are similar.

This section covers the complete CFD solution process including initialization, convergence criteria, and validation against experimental data. Hybrid initialization provides initial conditions for governing equations. Residuals should reach 1e-6-1e-7, but physical monitors (lift and drag coefficients) determine true convergence. Lift converges more reliably than drag due to Newtonian effects versus drag's sensitivity to surface roughness and separation. Expected accuracy of 10-20% deviation from experimental data is considered acceptable given CFD's inherent numerical approximations. Validation compares lift coefficients (NACA 4412 shows ~0.4 at zero degrees) and drag coefficients against sources like NASA Langley Research Center.

The study used Fluent U CFD software with RANS equations at 30 m/s free stream velocity (108 km/h). A mesh independence study varied cells from 2M to 32M, achieving 2% drag coefficient variation at 12M cells. Four turbulence models were tested; Realizable k-epsilon was selected for stability. Validation against 1/18th scale wind tunnel testing showed ~10% drag coefficient accuracy with good agreement on major pressure coefficient trends.

This tutorial demonstrates how to perform basic wind tunnel testing on aircraft models using Autodesk CFD software. The process involves creating a bounding box around the aircraft model in Fusion 360, exporting it as a STEP file, importing it into CFD, assigning materials (air for the fluid region and aluminum for solid components), setting boundary conditions (velocity at the inlet and zero pressure at the outlet), configuring simulation parameters including iterations and physics settings, and analyzing results using tools like the wall calculator for lift and drag forces, velocity planes for flow visualization, and smoke traces to observe airflow patterns around the aircraft.

This tutorial demonstrates the complete workflow for simulating a wind tunnel using ANSYS Fluent, covering geometry creation in Design Modeler with proper constraints and dimensions, mesh generation with appropriate element sizing, setup of boundary conditions including velocity inlet (10 m/s), pressure outlet, and wall boundaries, configuration of k-epsilon turbulence model with air as the working fluid, and post-processing techniques for visualizing static pressure, dynamic pressure, total pressure, and velocity magnitude contours along with centerline plots to analyze flow behavior through the wind tunnel section.

This tutorial demonstrates how to set up and run a CFD (Computational Fluid Dynamics) wind tunnel simulation using simulation software. The process involves importing a 3D model (STEP/STP file), creating a computational domain (air volume), assigning materials to the object and fluid, configuring boundary conditions (inlet velocity, outlet conditions), and setting simulation parameters like mesh count and temperature. After running the simulation, the key results are the drag force (resistance) and lift force (upward force) acting on the object, which are calculated by analyzing the wall forces. Higher mesh counts (400-500 or more) provide more detailed results but require longer computation times.

Wind tunnels employ various measurement techniques to analyze aerodynamic forces and flow patterns: 3D force balances measure forces and moments on objects using aerodynamic connection profiles; pressure taps with tubes and transducers capture surface static pressure distributions; smoke visualization traces airflow paths to identify separation points and vortex structures; Particle Image Velocimetry (PIV) uses tracer particles illuminated by lasers and tracked by high-speed cameras to calculate velocity vectors; and 3D measurement probes determine flow parameters like static pressure, total pressure, angle of attack, and velocity by measuring differential pressures at their multi-hole heads while tracking their 3D position.

Modern wind tunnel testing employs sophisticated instrumentation and measurement technologies to capture comprehensive aerodynamic data. Angle of attack measurement uses laser alignment tools taking multiple reference points on walls and floors to create geometric planes against which model position is precisely determined. Pressure measurement employs two complementary approaches: traditional methods drill thousands of small holes (approximately 40,000) with sharp lips to minimize flow disturbance, while modern pressure-sensitive paint technology offers an alternative covering entire model surfaces. This pink-colored paint changes color intensity in response to pressure variations, providing comprehensive coverage across complex geometries like wing tips and fuselage junctions while eliminating the need for individual hole drilling. The data collection system continuously captures information throughout each test run, with intelligent sampling techniques selectively capturing data only when specific criteria are met, improving data quality by filtering out transient variations and reducing scatter in final results.

A pitot static tube measures air velocity in a wind tunnel by comparing total pressure (stagnation pressure at the tube's tip) with static pressure (from ports along the tube wall); the device is installed on the test section wall with the open tip facing the inlet, connected to pressure transducers to calculate velocity distribution, and can be traversed vertically to measure velocity at multiple points across the test section.

The pitot-static tube is a fundamental aerodynamic instrument used to measure flow velocity in wind tunnels. It consists of a tube with a hemispherical front containing two sets of holes: one at the front aligned with the free stream flow, and multiple holes around the sides downstream. The front hole measures total pressure by capturing fluid that decelerates to zero velocity at the stagnation point, converting all dynamic pressure to static pressure. The side holes measure static pressure as fluid flows parallel to them. The relationship Total Pressure = Static Pressure + Dynamic Pressure enables calculation of flow velocity once both pressures are measured.

A Pitot-static tube measures dynamic pressure (also called velocity pressure). It consists of a Pitot tube (which measures stagnation pressure) and a static tube (which measures static pressure). The difference between stagnation pressure and static pressure gives the dynamic pressure, which can be used to calculate fluid velocity.

Compressible flow occurs when fluid speed approaches the speed of sound, causing significant density changes; the Mach number (ratio of flow speed to sound speed) determines flow regime: M < 0.3 is incompressible, M = 0.3-0.8 is subsonic compressible, M = 0.8-1.2 is transonic with shock waves, M > 1.2 is supersonic with shocks and expansion waves, and M > 5 is hypersonic with high temperatures and chemical reactions. Unlike incompressible flow analysis which only requires velocity and pressure, compressible flow analysis requires additional variables (density, temperature) and equations (conservation of energy, equation of state), and Bernoulli's equation no longer applies.

The Mach number (M = V/c) is the ratio of flow velocity to speed of sound, characterizing compressibility effects. Flow regimes are classified: subsonic (M < 1), sonic (M = 1), transonic (0.9 < M < 1.1), supersonic (M > 1), and hypersonic (M > 3). For air at 20°C, c ≈ 340 m/s, so 200 m/s gives M ≈ 0.59 (subsonic) while 600 m/s gives M ≈ 1.76 (supersonic). When objects travel supersonically, they generate shock waves forming a Mach cone behind them. Inside the cone, fluid is compressed with increased pressure; outside, fluid remains undisturbed. This phenomenon creates visible condensation clouds when shock waves pass through moist air.

The Mach number categorizes compressible flow regimes into four distinct categories: subsonic flow (M < 1) where flow velocity is below the speed of sound; transonic flow (around M = 1) characterized by mixed subsonic and supersonic regions; supersonic flow (M > 1) where objects move faster than sound waves, creating Mach cones and shock waves; and hypersonic flow (M > 5) where extreme temperatures cause gas ionization, dissociation, and communication blackouts. The transition between regimes is gradual rather than abrupt, with compressibility effects becoming significant when Mach exceeds 0.3.

The Mach number (M = U/a) classifies compressible flow regimes: (1) Subsonic (M < 1): Flow velocity less than sound speed, pressure waves propagate ahead of objects, smooth flow adjustment. (2) Sonic (M = 1): Critical transition point. (3) Supersonic (M > 1): Flow velocity exceeds sound speed, pressure waves cannot propagate ahead, shock waves form. (4) Hypersonic (M >> 1, typically >5): Molecular dissociation and ionization become important. (5) Transonic (M = 0.8-1.2): Flow can accelerate to sonic speed over the body, creating shock waves. The area-velocity relationship is opposite to incompressible flow: subsonic flow accelerates when area decreases, while supersonic flow decelerates when area decreases.

This section covers the theoretical foundation of aerodynamic flow regimes. The aerodynamic behavior of airfoils depends strongly on the Mach number, with compressibility effects becoming dominant as velocity increases from subsonic to transonic and supersonic regimes. In subsonic flow, the flow remains entirely subsonic with smooth pressure distribution and no shocks, and lift increases linearly with angle of attack. In transonic flow, local supersonic pockets develop on the suction side, terminated by normal shockwaves that cause boundary layer separation and significant drag increase at the drag divergence Mach number. In supersonic flow, the flow is mostly supersonic with bow shocks forming ahead of blunt leading edges and wave drag dominating, being highly sensitive to airfoil thickness and leading edge shape.