NACA 0012 Airfoil, Compressible Flow — ANSYS Fluent CFD SimulationDescriptionThis project simulates the flow over a NACA 0012 airfoil using ANSYS Fluent, with compressible flow as the central modeling theme. At the freestream conditions studied here, the air can no longer be treated as incompressible — density varies appreciably with pressure and temperature across the flow field — so the simulation is built around a compressible-flow formulation, making it a clear illustration of how that class of flow model is set up and solved. The airfoil is the cross-sectional shape of a lifting surface such as an aircraft wing, a wind-turbine blade, or a helicopter rotor, and its aerodynamic behavior depends strongly on its profile. The objective is to examine the airflow behavior and the pressure distribution around the airfoil and to study the resulting lift and drag forces. As the opening project of the Compressible Flow: Beginner CFD Training Package, it introduces the foundational external-aerodynamics case and the density-based solver setup that underpins the rest of the package.MethodologyThe geometry is defined by familiar parameters: the chord line, the leading and trailing edges, and the angle of attack — the angle between the chord and the oncoming flow direction. In this case the angle of attack is 5°, so the incoming velocity is resolved into a horizontal component of cos5° ≈ 0.996 and a vertical component of sin5° ≈ 0.087. The geometry is created in Design Modeler and meshed in ANSYS Meshing with a structured grid of 35,000 cells. Because the flow is compressible, a density-based solver is used — the appropriate choice when density variations are coupled tightly to the pressure and energy fields, as they are in high-speed aerodynamics. For compressible flow, the Mach number must be specified in the boundary conditions; it is the ratio of the flow speed to the local speed of sound (for reference, the speed of sound in air at 25 °C is about 343 m/s). Airfoil simulations of this kind require a far-field boundary condition with the Mach number prescribed for the surrounding flow, set here to 0.6 — firmly in the subsonic-but-compressible regime where compressibility effects are significant and cannot be neglected.AnalysisThe solution produces two-dimensional contours of pressure, velocity, temperature, density, and Mach number, together with streamlines around the profile. The results show the highest pressure at the leading edge, where the flow stagnates on direct contact with the airfoil, and the strongest pressure drop along the upper surface. This pressure difference between the upper and lower surfaces is what generates lift. The velocity field mirrors the pressure field exactly, as expected: regions of highest pressure coincide with the lowest velocity, and regions of lowest pressure with the highest velocity — the classic inverse relationship that underlies airfoil aerodynamics, here captured within a fully compressible treatment that also resolves the accompanying temperature and density variations. By the end of this project, you'll be able to set up a compressible external-flow simulation with a density-based solver, define far-field boundary conditions with a prescribed Mach number, and interpret the pressure, velocity, and Mach fields that characterize compressible airfoil aerodynamics.

DescriptionTransonic flow over an airfoil sits at the heart of compressible flow modeling, where local flow velocities approach and exceed the speed of sound, producing complex shock and expansion phenomena that govern aircraft aerodynamic performance. This CFD simulation uses ANSYS Fluent to model compressible airflow over a Naca 0012 airfoil in three dimensions, capturing transonic behavior at a Mach number of 0.7, an air temperature of 300 K, and a 2-degree angle of attack.MethodologyA two-dimensional airfoil geometry is constructed with multiple zone divisions in ANSYS Design Modeler and extruded to form a three-dimensional computational domain. The domain is discretized in ANSYS Meshing using a structured grid of over 1.5 million cells. The simulation employs a pressure-based solver configured for compressible flow, using the Coupled pressure-velocity coupling algorithm, ideal-gas behavior for air density, and the Sutherland model to capture temperature-dependent viscosity. The setup runs as a steady-state simulation appropriate for transonic flow conditions.Results AnalysisPost-processing examines velocity and pressure distributions around the airfoil to characterize the transonic flow field, including the effects of the 2-degree angle of attack on flow characteristics. The relationship between Mach number, local velocity, and pressure is analyzed to identify regions of compressibility effects and potential shock formation. These results provide insight into aerodynamic behavior in the transonic regime, relevant to aircraft design, aerospace research, and the broader study of compressible flow phenomena.

Wind Tunnel Drag Analysis — ANSYS Fluent CFD Simulation TrainingThis project models a wind tunnel and a specific body placed inside it using ANSYS Fluent, with the goal of investigating the drag force acting on the body.The geometry was created in ANSYS Design Modeler® and meshed in ANSYS Meshing® using an unstructured grid, for a total of 179,542 elements.BackgroundThe wind tunnel is one of the most widely used aerodynamic testing tools in use today. Among the many experiments it enables are tests of various structures, including airfoils, aircraft, and static bodies. These experiments are typically aimed at studying aerodynamic behavior and visualizing the flow lines around the test object.Wind tunnels can also be used for free-fall tests, examining how airflow affects a falling object. Because forces such as drag have a significant influence on a body's behavior, quantifying them is essential — and CFD provides an effective means of doing so.MethodologyA density-based (compressible flow) solver is used in this simulation, and the energy equation is activated accordingly to capture the compressible-flow physics.ResultsThe solution produces contours of velocity, pressure, temperature, and related quantities for a range of inlet Mach numbers. The velocity vectors reveal the formation of separation vortices behind the body. As a result of this separation, the flow turbulence in the wake is substantial — noticeably greater than in the rest of the computational domain.

Slat and Flap Devices Effects on an Aircraft Wing, ANSYS Fluent TrainingDescriptionThis project simulates the airflow around a 3-D aircraft wing fitted with a slat on the leading edge and a flap on the trailing edge, using ANSYS Fluent.This is a representative compressible-flow case. Near the wing the air moves fast enough that its density can no longer be treated as constant, so the flow must be solved with a density-based solver and the air modeled as an ideal gas whose density varies with pressure and temperature. Capturing these density changes — and the pressure field they produce over the wing — is exactly what the Compressible Flow model is built for.A flap is a small aerodynamic surface on the trailing edge of the wing used to increase lift. Lift is what holds the aircraft up against its weight, and it grows with speed. When the aircraft slows down — as during takeoff and landing — lift drops, so it must be recovered: the flap rotates about its hinge at the trailing edge and increases the effective area of the wing exposed to the airflow. A slat is the equivalent device on the leading edge; it rotates about its hinge to increase the wing's contact area with the air, and it also raises drag, which helps the aircraft land more slowly.The 3-D geometry was built in Design Modeler as a wing with a leading-edge slat and a trailing-edge flap. The model was meshed in ANSYS Meshing using an unstructured grid of 5,658,021 elements.Simulation MethodologyBecause the flow is compressible, the simulation uses a density-based solver with the ideal-gas law for density, and it is run as steady with gravity neglected. Turbulence is handled with the Spalart-Allmaras model, and the energy equation is enabled to resolve the temperature field.For boundary conditions, the inlet is a velocity inlet with components of 271.958 m/s along x and 40.79 m/s along y (a combined freestream of about 275 m/s at a small angle of attack) and a temperature of 305.5 K. The outlet is a pressure outlet at 0 Pa gauge, the wing, flap, and slat are walls with zero heat flux, and a symmetry plane bounds the domain. The flow and the turbulence transport variable are both discretized with a second-order upwind scheme, and the solution is initialized from the inlet conditions.Results & ConclusionAfter solving, two-dimensional contours of pressure, temperature, velocity, Mach number, and density were obtained on a plane cutting through the flow adjacent to the wing, along with path lines and velocity vectors on the same plane and a pressure contour over the wing surface.The contours clearly show the variation of velocity, density, and pressure around the wing. Comparing the pressure contour on the upper and lower surfaces reveals the pressure difference between them, which is the source of the lift force that counteracts the aircraft's weight.

DescriptionThis module uses ANSYS Fluent to simulate compressible flow through a 3D convergent-divergent nozzle, a fundamental problem in compressible flow modeling and high-speed propulsion applications. The simulation examines shock wave formation, Mach number evolution, and boundary layer behavior as flow accelerates from subsonic to supersonic conditions through the nozzle throat and divergent section.MethodologyThe nozzle geometry is analyzed under compressible flow conditions, resolving the transonic transition at the throat and the resulting supersonic flow field in the divergent section. Shock wave development is captured and visualized, along with its effects on pressure, temperature, and velocity. Boundary layer behavior along the nozzle walls is modeled under compressible conditions, including interactions between the boundary layer and shock waves, and flow separation regions in the divergent section are identified using CFD-based prediction methods.ConclusionResults include Mach number distributions showing the subsonic-to-supersonic transition, pressure and temperature profiles across shock regions (including deviations from ideal isentropic behavior due to friction and heat transfer), and flow separation zones affecting nozzle performance. Nozzle efficiency metrics such as thrust coefficient and specific impulse are calculated, and parametric studies support nozzle geometry optimization. These findings apply directly to rocket propulsion system design and supersonic wind tunnel development.

Supersonic Nozzle Flow Separation and Shock Wave — ANSYS Fluent CFD SimulationDescriptionWelcome to the Supersonic Nozzle Flow Separation and Shock Wave CFD Simulation module. This project explores supersonic nozzle flow using ANSYS Fluent, focusing on the intricate phenomena of flow separation and shock-wave formation. A supersonic nozzle accelerates gas from subsonic to supersonic speeds through a converging-diverging passage, but under off-design conditions the flow can separate from the walls and form shock waves — sharp discontinuities across which pressure, temperature, and Mach number change abruptly. Capturing these effects is central to the design of rocket engines, supersonic wind tunnels, and high-speed propulsion systems. Within the Compressible Flow: Beginner CFD Training Package, this project builds on the convergent-divergent nozzle case by focusing on the shock waves and boundary-layer separation that dominate supersonic nozzle behavior.MethodologyThe setup is built around the governing equations of compressible flow in the supersonic regime, with a turbulence model chosen to suit high-speed flow and its shock-turbulence interactions. The simulation captures the transition from subsonic to supersonic flow through the converging-diverging passage, resolving both normal shock waves and the oblique shock structures that form in an overexpanded nozzle and interact with the walls. The boundary-layer development is modeled to reveal where the flow separates under the adverse pressure gradient, and the setup supports varying the back pressure to study how the nozzle flow adapts across overexpanded and underexpanded regimes.AnalysisThe results provide Mach number contours that reveal the flow acceleration and shock formation along the nozzle, together with the pressure and temperature jumps across the shocks — the pressure discontinuities and the temperature rise from shock compression that matter for material selection and thermal management. From these you can identify the separation point under different pressure ratios, quantify nozzle performance through the thrust coefficient and efficiency, and visualize complex shock structures such as shock diamonds. By the end of this project, you'll be able to set up a supersonic nozzle simulation with a compressible solver, capture normal and oblique shock waves and boundary-layer separation, study the effect of back pressure on the flow, and interpret the Mach, pressure, and temperature fields that govern rocket-nozzle and high-speed propulsion design.

Steam Ejector — ANSYS Fluent CFD SimulationDescriptionThis project presents a CFD simulation of a steam ejector — a mechanical device with no moving parts that uses a primary (motive) steam jet to suck in and mix with a secondary fluid. Ejectors perform two essential jobs: creating vacuum for suction and mixing two fluid streams, and they do it by continuously converting between kinetic and pressure energy as the flow passes through a convergent-divergent nozzle. In this project, you'll model water vapor as the motive fluid driving the suction of a secondary fluid, watching the flow accelerate beyond the speed of sound and observing how the vacuum-driven suction physically arises. Within the Compressible Flow: Beginner CFD Training Package, this project applies convergent-divergent nozzle physics to a real industrial device, showing how supersonic internal flow generates suction and mixing.MethodologyThe 2D convergent-divergent (de Laval) nozzle ejector geometry is designed in Design Modeler and meshed in ANSYS Meshing with an efficient structured grid of roughly 52,000 elements suited to internal compressible flow. Because supersonic flow is fundamentally compressible — with density varying strongly with pressure — the density-based solver is used, the correct choice for this class of problem, and the Mach number governs the behavior inside the device. The setup handles the pressure difference between the primary and secondary inlets that drives the suction phenomenon, so the motive steam jet entrains and mixes with the secondary fluid as it accelerates through the nozzle.AnalysisPost-processing focuses on pressure, velocity, and Mach number contours, tracing where the flow goes subsonic, sonic, and supersonic through the device. From these you can follow how the motive and secondary streams mix and compress downstream of the nozzle throat, and how the vacuum-driven suction arises from the energy conversion in the convergent-divergent passage. Ejectors appear throughout refrigeration, vacuum systems, desalination, chemical processing, and power plants, and the skills built here carry directly into nozzles, diffusers, supersonic airfoils, and any flow where Mach number matters. By the end of this project, you'll be able to set up an internal compressible-flow simulation with the density-based solver, handle the inlet pressure difference that drives ejector suction, and interpret the pressure, velocity, and Mach fields that trace the subsonic-to-supersonic behavior inside the device.

DescriptionThis simulation examines flow through a rocket engine nozzle in ANSYS Fluent, a compressible flow problem central to propulsion design. A rocket nozzle works by converting the internal energy of hot propellant gases into directed kinetic energy, producing a high-velocity exhaust jet; the gases enter subsonic, are forced to accelerate as the passage narrows toward the throat, and reach sonic velocity exactly at that minimum cross-section. Beyond the throat, the diverging section allows the gas to keep expanding and accelerating, pushing the flow to supersonic exit speeds. The geometry is built as a 2D convergent-divergent nozzle in Design Modeler, capturing the interior flow path through the throat and diverging sections, and meshed in ANSYS Meshing with an unstructured grid of 69,342 cells.MethodologyBecause the flow is compressible and spans subsonic-to-supersonic regimes across the nozzle, the density-based solver is used, with gas density computed from the ideal gas law rather than treated as constant. The nozzle inlet is set to a gauge pressure of 2,268,000 Pa, and the outlet to 39,365 Pa, a large pressure ratio that drives the expansion and acceleration through the throat.AnalysisThe solution produces 2D contours of temperature, pressure, velocity, density, and enthalpy, along with streamlines through the nozzle. These fields confirm the expected compressible-flow behavior: the gas accelerates sharply as it passes through the throat, and this acceleration is accompanied by a corresponding pressure drop as the cross-sectional area shrinks, consistent with the physics of convergent-divergent nozzle flow.

DescriptionThis project simulates an aerospike using ANSYS Fluent. The goal is to model a nose cone fitted with an aerospike and to study how it reshapes the shock wave — pushing it away from the nose cone — and thereby reduces drag. Because the vehicle travels at supersonic speed, the flow is fully compressible, and the analysis centers on resolving the shock structure that forms ahead of the body — a defining feature of compressible flow.A drag-reducing aerospike is a device used to lower the forebody pressure drag of blunt bodies at supersonic speeds. The aerospike creates a detached shock ahead of the body, and a zone of recirculating flow forms between the shock and the forebody. This recirculation zone acts like a more streamlined forebody profile, which reduces the aerodynamic drag.The geometry was created in ANSYS Design Modeler, and meshing was performed in ANSYS Meshing using an unstructured grid with a total of 153,987 cells. The figure below shows an overview of the mesh.MethodologyThe simulation uses the k-omega SST turbulence model, with the flow set to Mach 2. At this speed the air must be treated as a compressible medium, so the solver captures the density variations and shock waves that govern the aerodynamics of the aerospike.ConclusionOn completion of the solution, two-dimensional velocity contours and an animation were obtained. The results show that the aerospike produces a much thinner and weaker oblique shock, positioned well away from the cone. Because the pressure gradient of the primary shock wave now sits far from the body, the drag is reduced.A small separation region also forms on the cone itself, and the reattachment of the separated flow together with the oblique shock occurs at some distance from the main body — which reduces the aerothermodynamic efficiency. Overall, the simulation illustrates a classic compressible-flow phenomenon: how modifying the shock structure ahead of a supersonic body can achieve meaningful drag reduction on a blunt forebody.

DescriptionThe Lockheed Martin F-35 Lightning II is an American family of single-seat, single-engine, all-weather stealth multirole combat aircraft, also capable of electronic warfare and intelligence, surveillance, and reconnaissance missions.Notably, the F-35 can reach speeds of around 500 m/s, placing the surrounding airflow firmly in the supersonic regime. Since full-scale wind tunnel experiments at these conditions are costly in both time and money, CFD solvers are frequently used for initial evaluation. This project studies the supersonic, compressible flow around an F-35 aircraft. The geometry consists of a 20 m F-35 positioned inside a 150 m wind tunnel.The mesh contains 7,182,542 elements. In terms of quality, a maximum skewness of 0.79 with an average of 0.22 is satisfactory for this problem. To resolve the boundary layer accurately, 25 prism layers were added adjacent to both the wind tunnel walls and the aircraft body. The mesh was generated in ANSYS Meshing and subsequently converted to a polyhedral mesh within ANSYS Fluent, reducing the count to 1,845,364 elements while preserving the same quality. As with any numerical study, the first step in the modeling was the creation of the CAD geometry.MethodologyAir is treated as a compressible ideal gas, and a Mach number of 2.0 is reached at the maximum speed of 544 m/s. Solving this problem requires the flow equations to be handled in their differential form.A non-isothermal, compressible ideal-gas condition was assumed inside the wind tunnel; consequently, the energy equation was solved alongside the flow and turbulence equations. The governing mass and momentum equations are written in their standard conservative form for compressible flow.ConclusionThe drag force and the shock profiles were obtained over the course of the study. After the solution converged, the results were examined through post-processing. As a check on convergence, the drag value was monitored throughout the solution iterations: the solution was deemed converged once the drag force settled to a constant value and the residuals dropped below 10⁻⁶.The results are then presented for the pressure and velocity fields. The shock profile is visible in both the pressure and Mach number contours, while the velocity field is shown through both contours and streamlines to give deeper insight into the flow. The temperature gradient and its variation across different locations are also presented, since the temperature rise is an important factor in compressible aerodynamic calculations. Finally, the drag force was calculated at 181.66 kN, a reasonable value for a 20 m aircraft with the stated specifications.