Circulation control improvements to rotor lift asymmetry due to reverse flow

The flows about 2-D and 3-D bluff trailing edge circulation control (CC) airfoils are computed using steady Reynolds Averaged Navier-Stokes (RANS) methods. The 2-D foil is the NCCR 1510-7067 elliptical CC airfoil with circular and logarithmic spiral trailing edge geometries. The free stream Reynolds number, based on chord, is 5.45 × 10 5 , with a free stream Mach number of 0.12. For the circular trailing edge the slot height, blowing rate and angle of attack are varied, while for the logarithmic spiral only the blowing rate is varied. The 3-D foil is a semi-span wing with an elliptical cross section. It is run with a chord-based Reynolds number of 2 × 10 6 and two blowing rates. The 2-D flows are computed using the compressible, segregated solver, Fluent. 2-D results show that the full-Reynolds stress turbulence model (FRSM) predicts the correct jet detachment behavior for the circular trailing edge although the integrated lift forces are consistently underpredicted. The coanda jet detachment point for the logarithmic spiral trailing edge is predicted correctly for a lower blowing rate, but as blowing rate increases, the jet does not detach until it has wrapped around to the pressure side. We show additional 2-D results using mesh refinement via grid adaption and isotropic eddy viscosity turbulence models. The 3-D simulations use the incompressible segregated Fluent solver applying the k −ω SST turbulence model. Results show a slight attachment of the the coanda jet on the pressure side, but the results are generally encouraging.

Lift enhancement based on virtual aerodynamic shape using a dual synthetic jet actuator

Short-takeoff marine vehicle with circulation flow control of wing-in-ground effect

In order to study the influence of elastic forward-swept wing (FSW) with single control surface, the computational fluid dynamics/computational structural dynamics (CFD/CSD) loose coupling static aero elastic numerical calculation method was adopted for numerical simulation. The effects of the elastic FSW with leading- or trailing-edge control surface on aero elastic characteristics were calculated and analysed under the condition of high subsonic speed. The result shows that, the deflection of every single control surface could change the aero elastic characteristics of elastic FSW greatly. Compared with the baseline model, when leading-edge control surface deflected up, under the condition of small angles of attack, the aerodynamic characteristics was poor, but the bending and torsional deformation decreased. Under the condition of moderate angles of attack, the aerodynamic characteristics was improved, but bending and torsional deformation increased; When leading-edge control surface deflected down, the aerodynamic characteristics was improved, the bending and torsional deformation decreased/increased under the condition of small/moderate angles of attack. Compared with the baseline model, when trailing-edge control surface deflected down, the aerodynamic characteristics was improved. The bending and torsional deformation increased under the condition of small angles of attack. The bending deformation increased under the condition of small angles of attack, but torsional deformation decreases under the condition of moderate angles of attack. So, for the elastic FSW, the deflection of trailing-edge control surface play a more important role on the improvement of aerodynamic and elastic deformation characteristics.

Experimental testing was performed on a circulation controlled airfoil with upper and lower trailing edge blowing slots, controlled by span wise pneumatic valves. The augmented blade was designed for application to a circulation controlled vertical axis wind turbine. The design is based upon a conventional NACA0018 shape, replacing the sharp trailing edge with a rounded Coanda surface and blowing slots. A scale model with a chord of 8 inches and span of 16.5 inches was created using an ABS plastic rapid prototyping machine. In the past, circulation control wind tunnel models have been constructed with a separate blowing slot and trailing edge using conventional machining methods. The slot must be tediously aligned along the span for a consistent height which ultimately affects the uniformity and performance of the circulation control jet in combination with the flow rate. The rapid prototyping machine eased fabrication as a modular trailing edge section was printed which includes the Coanda surface, blowing slot, and diffuser all in one piece. Pressure taps were integrated by the prototyping machine into both the printed skin and trailing edge module. This method left additional space inside the model for circulation control valving components and eliminated the need for machining pressure ports. This paper will outline the model building procedures, wind tunnel test rig, and experimental results. Aerodynamic forces were determined by both load cells and surface pressure measurements; the agreement between the two methods will be analyzed and addressed. Test conditions include various angles of attack (±20°) at Cμ = 0, 0.02, 0.06, and 0.10; the test Reynolds number was kept constant at 300K. The results indicate that the blade performed at ΔCl/Cμ near 30 for Cμ = 0.02.

During spacecraft reentry, the aerodynamic characteristics of the reentry capsule with a sizeable obtuse head inverted cone (OBHIC) in a hypersonic state are the key factors in the design of controlled reentry trajectory, the design of trim angle of attack (AoA), and the accuracy of fall point of capsules. This paper establishes a numerical simulation algorithm of a hypersonic chemical nonequilibrium flow field by considering high-temperature real gas effects. The reentry capsule’s aerodynamic and flow field characteristics are predicted and analyzed. The component transport equations for chemical reactions are added to the Navier-Stokes equations, enabling precise modeling of gas molecule vibrational excitation and ionization processes. A half-model multi-block structured mesh is employed to construct the aerodynamic profile of the reentry capsule with OBHIC. The aerodynamic characteristics and pressure distribution of the reentry capsule were calculated and analyzed under the typical conditions of flight altitude of 40~80 km, flight velocity of 6~27 Ma, and AoA of 15°~24°. The results of the aerodynamic characteristics prediction show that the aerodynamic characteristics of the reentry capsule are little affected by the incoming flow velocity in a hypersonic state but significantly affected by the AoA. As the AoA increases, the lift coefficient CL increases linearly, the drag coefficient CD decreases linearly, the lift-to-drag ratio K rises linearly, and K increases from 0.19 to 0.29. In its hypersonic state, the reentry capsule exhibits a low lift-to-drag ratio, characterized by high drag and low lift coefficients. The results of this paper play an essential role in the reentry flight dynamics and reentry trajectory design of the reentry capsule. Moreover, it can further ensure the safety and reliability of the reentry flight of spacecraft.

Analytical prediction of trajectories of ice pieces after release in an airstream

This study investigates the influence of the Strouhal number on the thrust and propulsive efficiency of two-dimensional tandem elliptical flapping airfoils through numerical simulations validated by in-house experiments at a Reynolds number of 5000. The front and rear airfoils, identical in shape and size, are aligned with the freestream, with the rear airfoil positioned one chord length behind the front. The Strouhal number was varied from 0.2 to 0.7, while the geometric effective angle of attack amplitudes at the pivot points were set to 10, 15, and 20 degrees. Results show that the rear airfoil has minimal effect on the time-averaged thrust of the front airfoil, but the wake of the front airfoil strongly influences the flow field and thrust of the rear one. The thrust produced by the rear airfoil increases with the Strouhal number up to a critical value of 0.4, then decreases, with larger angles of attack producing higher thrust. The growth of the leading-edge vortex on the rear airfoil is controlled by the residual vortex beneath the front airfoil and the shear layer shed from the upper surface of the front airfoil. The residual vortex enhances the local flow angle and increases thrust, while the shear layer intermittently reduces it. When the shear layer passes over the rear airfoil, the local flow angle rises sharply, leading to rapid vortex growth and greater thrust, particularly at a 15-degree angle of attack. These results demonstrate the aerodynamic advantages of tandem flapping configurations under optimized unsteady flow interactions.

The reverse-flow theorem is extended to the case of a wing mounted on a cylindrical body of any cross-sectional shape. Also, two TrefTtz-plane theorems relating the total lift and rolling moment on such a body to the corresponding values for the wings are derived. Based on these results, some further general relationships between integrated lifts and moments on the wing, the body, and the combination are obtained. I t is shown that the lift distribution on the wing-body combination in reverse flow is the influence function for lift of the combination. Further, it is shown that the influence function for rolling moment is the lift distribution over the combination rolling at constant angular velocity in reverse flow. I t is also shown tha t the lift induced on the wings by a unit body angle of attack is equal to the lift induced on the body by a unit wing angle of attack in reverse flow. In lifting-line theory, slender-body theory, or for any wings having a span wise axis of symmetry (such as rectangular or diamond wings), the need for consideration of reverse flow is eliminated, and the results assume a particularly simple form. The results are valid for subsonic or supersonic flow within the limitations of linearized theory.

This paper puts forward an aerodynamic parameter estimation method which combines the Wavelet Neural Network(WNN) with Modified Particle Swarm Optimization(MPSO) technique. This method directly and accurately constructs the relationship among flight parameters and aircraft aerodynamic parameters. Preliminary aerodynamic parameters and derivatives are derived from wind tunnel data. With strong nonlinear mapping ability, WNN is used to construct the relationship among Mach number, angle of attack, and rudder with longitudinal aerodynamic parameters. Then, the MPSO is used to estimate aerodynamic parameters based on the test-flight data. And WNN is retrained to amend the relationship among Mach number, angle of attack, and rudder with longitudinal aerodynamic parameters. Simulation verification indicates that MPSO has better estimation accuracy than Maximum Likelihood(ML) method. Comparison results of simulation experiments and flight-test data of a tactical missile show that simulated data based on estimated parameters matches with the flight-test data, which prove the effectiveness and validity of the trained WNN.

PurposeThe purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing.Design/methodology/approachSimulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference.FindingsIt was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs.Practical implicationsFor accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation.Originality/valueThe presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.

Swirl and buoyancy effects on mixing performance of baffle-plate-type miniature confined multijet

This study investigates the influence of the angle of attack (AOA) of the NACA 2412 airfoil on aircraft fuel efficiency across different flight speeds and determines the optimal AOA for maximum performance. Using a quantitative modeling approach implemented in Python, the research establishes an aerodynamic–fuel consumption relationship under steady, level flight conditions (lift = weight, thrust = drag). The model draws on validated aerodynamic parameters from NASA and FAA reports to ensure reliability. Simulation results show that at low speed (200 km/h), fuel efficiency remains relatively insensitive to AOA variation, indicating greater operational tolerance during takeoff and climb. However, at high speed (800 km/h), even small deviations from the optimal AOA (approximately −2.2°) lead to a rapid, nonlinear increase in drag and fuel consumption. These findings confirm that precise AOA control is crucial for high-speed cruise efficiency. The study contributes to the limited body of research linking aerodynamic parameters directly to energy use, providing a replicable framework for integrating AOA optimization into flight trajectory planning, autonomous control systems, and sustainable aviation strategies.

In this study, the dynamic stall evolutions were investigated using particle image velocimetry (PIV) in a water channel with Reynolds number Re = 4.5 × 103 based on the chord length. The airfoil pitching waveform was performed under the condition calculated from the angle of attack histogram of a vertical axis wind turbine (VAWT). Using PIV, the instantaneous vorticity contours and streamlines can be revealed. Based on the formation of the leading edge vortex, the stall angle can be explored at reduced frequency k = 0.09, 0.18, and 0.27. It was found that the stall angle was delayed from the angle of attack α = 16° to α = 30° as reduced frequency increased from k = 0.09 to 0.27. The hysteresis effect of stall angle delay was more pronounced for high reduced frequency. Moreover, the freestream turbulence effect on the pitching airfoil was investigated with turbulence intensity TI = 0.5 and 6.9 %. As found, the stall angles were postponed to higher angles of attack for the high turbulence intensity. The phase difference between TI = 0.5 and 6.9 % were ∆α = 8°, 4°, and 4° for k = 0.09, 0.18, and 0.27, respectively. For TI = 6.9 %, enhanced turbulence mixing reduces the velocity deficit (u/U < 1) and flow reversal (u/U < 0). In addition, the maximum velocity is reduced from u/U = 1.8 to 1.2 and the S-shaped velocity profile is diminished or weakened for TI = 6.9 %. Thus, the dynamic stall is further delayed to the downstroke. The circulation values increase rapidly to maximum and then drop quickly after dynamic stall for k = 0.18 and 0.27.

: A circulation control uncambered elliptic airfoil section with a thickness-to-chord ratio of 0.20 was tested subsonically to determine its aerodynamic characteristics. Lift coefficients up to 5 were produced at momentum coefficients, of 0.24. The initially high unblown drag coefficients, characteristic of bluff trailing edge airfoils, were greatly reduced at low values of momentum coefficient. It was therefore possible to produce equivalent lift-to-drag ratios in excess of 30 when Cl = 1.0. The ability to produce high lift coefficients essentially independent of angle of attack is indicated by the results of this investigation.