Deg Angle Of Attack Research Articles

Interactive Airfoil Optimization Using Parsec Parametrization and Adjoint Method

In the development of interactive aerodynamic optimization tools, the need to reduce the computational complexity of flow calculations has arisen. Computational complexity can be reduced by estimating the flow variables using machine learning, but that approach has a number of hindrances. Avoiding these hindrances through lowering the computational complexity by stating the assumptions of inviscid incompressible potential flow is the focus of this article. The assumptions used restrict the applicability of this approach to only specific cases, but in engineering practice, these cases are quite widespread. The assumptions allowed the coupling of the adjoint method with parsec parametrization and the panel method, yielding a highly computationally efficient and robust tool for optimizing an airfoil’s lift coefficient (Cy). The optimization of the NREL S809 airfoil was carried out, and the results were verified using the Xfoil 6.99 software. The Xfoil verification showed that by making minimal changes to the airfoil’s shape, the Cy and lift-to-drag ratios were significantly improved. The improvement magnitude was over 94% for a 0 deg angle of attack (AoA) and over 16% for 6.2 deg AoA. This indicates an improvement in performance that is similar to that of some genetic algorithms, but with computational costs that are many orders of magnitude lower.

Resolvent model for aeroacoustics of trailing edge noise

This study presents a physics-based, low-order model for the trailing edge (TE) noise generated by an airfoil at low angle of attack. The approach employs incompressible resolvent analysis of the mean flow to extract relevant spanwise-coherent structures in the transitional boundary layer and near wake. These structures are integrated into Curle’s solution to Lighthill’s acoustic analogy to obtain the scattered acoustic field. The model has the advantage of predicting surface pressure fluctuations from first principles, avoiding reliance on empirical models, but with a free amplitude set by simulation data. The model is evaluated for the transitional flow (Re=5e4\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\ ext {Re} = 5e4$$\\end{document}) around a NACA0012 airfoil at 3 deg angle of attack, which features TE noise with multiple tones. The mean flow is obtained from a compressible large eddy simulation, and spectral proper orthogonal decomposition (SPOD) is employed to extract the main hydrodynamic and acoustic features of the flow. Comparisons between resolvent and SPOD demonstrate that the physics-based model accurately captures the leading coherent structures at the main tones’ frequencies, resulting in a good agreement of the reconstructed acoustic power with that of the SPOD (within 4 dB). Discrepancies are observed at high frequencies, likely linked to nonlinearities that are not considered in the resolvent analysis. The model’s directivity aligns well with the data at low Helmholtz numbers, but it fails at high frequencies where the back-scattered pressure plays a significant role in directivity. This modeling approach opens the way for efficient optimization of airfoil shapes in combination with low-fidelity mean flow solvers to reduce TE noise.Graphical abstract

Numerical Simulation of a Y-Shaped Diverterless Supersonic Inlet

Three-dimensional (3-D) numerical simulations of a Y-shaped body-integrated diverterless supersonic inlet (DSI) have been investigated. A diverterless supersonic inlet is an external supersonic flow compression that uses a 3-D geometry of a bump for both supersonic flow compression and boundary-layer diversion. The simulations are performed in the presence of a typical body comprising an elliptical nose in order to model a more realistic approaching flow to the inlet entrance. All simulations were performed at a freestream Mach number of , the design Mach number, and at a 0 deg angle of attack (AOA) and 0 deg angle of sideslip (AOS). The Navier–Stokes relations along with shear stress transport equations are used to simulate turbulent flow on a multiblock structural grid. The present numerical simulation results show a good agreement with the wind-tunnel test data for the same DSI model. Moreover, the results show that the present numerical method has an acceptable accuracy in modeling a fully 3-D flow over the nose, along the supersonic compression region, and inside the inlet diffusers. The streamlines passed over the DSI supersonic compression clearly showed that the bump was successfully efficient in diverting the approaching boundary layer.

Large Eddy Simulation of Transitional Separated Flow Over a Low Reynolds Number Cambered Airfoil

Abstract The accurate simulation of the aerodynamic behavior of low Reynolds number (Re) cambered airfoils requires the ability to capture the transitional separated boundary layer (BL) that occurs naturally on the surface of the airfoil. In this study, simulations are performed using a modern cambered airfoil designed for use in low Re applications, which are an advancement from previous studies using flat plate geometries or symmetric NACA airfoils. The cambered SD 7037 airfoil is simulated using wall-resolved large eddy simulation (LES) at a modest Re of 4.1×104 and at 1 deg, 5 deg, and 7 deg angles of attack (AOAs), with results validated against experimental data. Simulated predictions of pressure and skin friction coefficients clearly capture the correct location of the laminar separated bubble (LSB) which forms during the natural BL transition process. Sensitivity to elevated inflow turbulence is found to cause early BL reattachment at higher AOAs without impacting the location of BL separation. An integral BL analysis verifies the accuracy of the simulated velocity profiles against experimental values. The scale of horseshoe structures visualized in the transitional BL is larger in comparison to airfoil chord length than what is seen in previous simulations at Re of the order of 105, which highlights the importance of investigating cambered airfoils at a modest Re.

Effect of angle of attack on the performance of the supersonic intake for High Mach Integrated Control Experiment (HIMICO)

Experimental, numerical and theoretical investigations are carried out to grasp the effect of angle of attack (AoA) on the performance of the supersonic intake for High-Mach Integrated Control Experiment, HIMICO. The results of theoretical analysis using simple shock wave (TFD) suggests that mass capture ratio (MCR) and total pressure recovery (TPR) are increased with positive angle of attack. In fact, comparing the result of 2.5 deg AoA with that of −2.5 deg AoA gained by CFD, MCR and TPR become 21pt and 5pt larger when the nozzle throat height (HNT) is 12.5 mm. The main factors of these phenomena are the increased projected area of the intake entrance and the change of the oblique shock system. On the other hand, TFD incorrectly estimates HNT and TPR in critical state because the pseudo-shock wave is generated instead of the terminal normal shock wave considered in TFD. So the characteristics of the pseudo-shock wave are investigated by CFD. As a result, it was found that the total pressure loss due to the pseudo-shock wave is larger than that of the normal shock wave, and that the total pressure loss becomes smaller with positive AoA due to the decrease in the Mach number at the intake throat. Using the CFD results, we created a model to estimate TPR and HNT in the critical state from the relational equation between the intake throat Mach number and the total pressure loss due to the pseudo-shock wave.

Bluntness-Dependent Hypersonic Boundary-Layer Modes’ Excitation

Boundary-layer disturbances on a cone in a hypersonic flow are investigated by heat flux sensors. The cone has a half-angle of 10.9 deg and is situated with a 0 deg angle of attack in a Mach 10 region of a conical nozzle in a shock-tunnel facility. There are three atomic layer thermopile heat flux sensors mounted at different chordwise positions in order to investigate the receptivity to freestream disturbances and the emergence of unstable modes of the boundary layer, that is, Mack first and second modes. The cone’s nose is exchangeable, and there are three differently sized radii to investigate the bluntness effect on transition with respect to entropy/boundary-layer displacement. It is found that each nose type tested shows a different set of amplified spectral bands in the receptivity region of the first two sensors, as well as at the third sensor’s location downstream of the critical point. There is a unique wavelength, scaling with boundary-layer height, that is most receptive to disturbances. For the two sharper cones, this excitation lies in the range of Mack second modes, which are observed in the unstable region. For the blunter cone, this excitation lies near the first mode frequency; as a result, only first modes are amplified and detected at sensor 3, although linear stability theory predicts higher growth rates for second modes.

A New Turbulent Viscosity Correction Model With URANS Solver for Unsteady Turbulent Cavitation Flow Computations

Abstract Due to the ignorance of the effect of the water–vapor interface on the cavitation flow field, the standard k–ε turbulence model (ST model) may overestimate the turbulent viscosity. It is unable to simulate cavitation shedding, especially at small attack angles of a hydrofoil. In the present investigation, a turbulent viscosity correction model is proposed to dampen the turbulent viscosity at the water–vapor interface. Cavitation flow around a NACA0009 truncated hydrofoil with a 2.5 deg angle of attack is used to demonstrate the effect of correction. The results show that the interface effect-based correction model (IE model) can both predict the pressure distribution on the suction surface of the hydrofoil with experimental data and the re-entrance jet in the leading-edge cavitation shedding. The region of the IE model influenced concentrates on the water–vapor interface and intensifies the vortex strength, which directly enhances the formation of a horseshoe vortex. The reduction of turbulent viscosity by the IE model reduces the resistance to the development of a re-entrance jet. The shear stress plays an important role in the shedding of the attached cavity bubble. The increase of shear force in the leading-edge cavitation occurs with the re-entrance of water and the main shear flow concentrates on the middle of the cavity bubble. This paper therefore presents a new method of numerical simulation of cavitation flow in engineering applications.

Symmetric and Asymmetric Performance Investigation of a Diverterless Supersonic Inlet

A series of wind-tunnel tests were conducted on a Y-shaped diverterless supersonic inlet (DSI) in order to investigate its performance at various operating conditions. The present DSI was designed for a freestream Mach number of . The experiments were performed at supercritical, critical, and subcritical operating conditions and at 0 deg angle of attack and angle of side slip. The results showed that the present DSI had an acceptable performance at its design condition. A symmetric supersonic flow pattern was observed at both supercritical and critical operating conditions. At low subcritical operating conditions, the supersonic flow pattern remained symmetric too. However, at high subcritical operating conditions, an asymmetric supersonic flow pattern was observed at both 0 deg angle of attack and sideslip angle. Moreover, an asymmetric pressure ratio distribution inside both diffusers, left and right, were observed. This asymmetric flow behavior in a perfectly symmetric Y-shaped DSI inlet seems to have not been noticed by the previous researchers, but if it has, the result has not been published yet. The performance curve of the current DSI presented at both symmetric and asymmetric flow patterns reveal a significant reduction in its performance characteristics at high subcritical operating conditions that is believed to be caused by the asymmetric flow patterns. In addition, symmetric and asymmetric secondary normal shock- wave patterns were further observed during both supercritical and high subcritical operating conditions. It is observed that the flow asymmetry in this inlet happened when the inlet exit area had blocked more than 50% of the aerodynamic interface plane.

The aerodynamics of flying snake airfoils in tandem configuration.

Flying snakes flatten their body to form a roughly triangular cross-sectional shape, enabling lift production and horizontal acceleration. While gliding, they also assume an S-shaped posture, which could promote aerodynamic interactions between the fore and the aft body. Such interactions have been studied experimentally; however, very coarse models of the snake's cross-sectional shape were used, and the effects were measured only for the downstream model. In this study, the aerodynamic interactions resulting from the snake's posture were approximated using two-dimensional anatomically accurate airfoils positioned in tandem to mimic the snake's geometry during flight. Load cells were used to measure the lift and drag forces, and flow field data were obtained using digital particle image velocimetry (DPIV). The results showed a strong dependence of the aerodynamic performance on the tandem arrangement, with the lift coefficients being generally more influenced than the drag coefficients. Flow field data revealed that the tandem arrangement modified the separated flow and the wake size, and enhanced the lift in cases in which the wake vortices formed closer to the models, producing suction on the dorsal surface. The downforce created by the flow separation from the ventral surface of the models at 0deg angle of attack was another significant factor contributing to lift production. A number of cases showing large variations of aerodynamic performance included configurations close to the most probable posture of airborne flying snakes, suggesting that small postural variations could be used to control the glide trajectory.

Experimental Investigation of Transverse and Vortex Gust Encounters at Low Reynolds Numbers

This paper aims to study the lift response of transverse and vortex gust encounters. A sine-squared transverse gust was created in a water towing tank using a water jet, whereas a vortex gust was created in a water channel by clockwise half-rotation of an upstream gust generator plate. Forces and flowfields were acquired for both gust types. The effects of the gust ratio and vertical positioning of the gust generator plate were investigated for the transverse gust and the vortex gust encounters, respectively. For a flat plate at a 0 deg angle of attack, changing the gust ratio for the transverse gust was found to have a major effect on the peak lift and a minor effect on the recovery phase; whereas changing the vertical positioning of the gust generator plate for the vortex gust was found to cause relatively minor effects at almost every point in the lift response. Both gust encounters resulted in large transients in the lift force, and the unsteady forcing was found to depend on the circulation shed from the leading edge of the wing. The lift force experienced and the circulation shed showed a steeper increase for the transverse gust encounters despite the same gust ratio of one. Average velocity variations in time were obtained from the flowfield measurements along an upstream line for both gust types, and variations in the gust-induced angle of attack were computed. The Helmbold equation was then used to predict the lift coefficients based on the calculated gust-induced angle of attack. The trends in the measured and predicted lift coefficients were found to compare well with each other, but the values were found to agree well only when the gust-induced angle of attack was low.

Focused Laser Differential Interferometry with Contoured Tunnel Windows

A unique focused laser differential interferometry (FLDI) configuration designed to work with contoured windows was developed for use with the Boeing/U.S. Air Force Office of Scientific Research Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The modified FLDI uses cylindrical lenses to help shape the beam profile together with the curved windows. The design was informed by a ray-tracing simulation of the apparatus. The spatial sensitivity of the BAM6QT FLDI was tested using density-fluctuation measurements of a small-diameter air jet and compared to a traditional FLDI. To validate the BAM6QT FLDI’s performance with the tunnel, second-mode fluctuation measurements were made on a 7 deg half-angle cone at a 0 deg angle of attack. Both conventional-noise and quiet-flow measurements were made with good peak frequency agreement between the FLDI and associated surface-pressure measurements. The density-fluctuation amplitudes increase to a peak and then decrease with increasing distance from the model surface, agreeing with computational predictions. Cross correlation of the FLDI and surface pressure sensor signals resulted in an estimated disturbance velocity in the downstream direction, as well as the phase shift off the model surface.

Aerodynamic Analysis of Hybrid Drone

Drones are Unmanned Aerial Vehicles (UAV’s) or Remotely Piloted Vehicles (RPV’s) whose application ranges from photography to defence. There are two major classifications of drone, one, a fixed wing; and another, a quadcopter. A hybrid drone is a kind of drone which has a mix of features of the two kinds of drone. It has quad-rotors for Vertical Take-Off Landing (VTOL), and a fixed wing for cruise. Aerodynamic analysis for a hybrid drone is done using Computational Fluid Dynamics (CFD) as a tool. A virtual wind tunnel test is done by means of this analysis. It is done in order to examine the various technical edges of the hybrid drone over the rest of the types. Firstly, the main entity of the drone, the wing, of 3 different profiles(aero-foils), is analysed for better lift coefficient as it is the major contributor for the thrust during cruise (for 0 deg angle of attack). The velocity contours are viewed to interpret the aerodynamic advantage of each wing. Then the whole vehicle is analysed after inculcating the wing. The various aerodynamic parameters such as the lift force, lift coefficient, drag force and drag coefficient of the hybrid drone are determined from the analysis. A hybrid drone with optimum lift has been designed from the inference.

Investigation of Numerical Hydrodynamic Performance of Deformable Hydrofoil (Applied on Blade Propeller)

The hydrofoil is a hydro-lifting surface that significantly contributes to marine transportation such as a boat, ship, and submarine for its movement and maneuverability. The existing hydrofoils are in fixed-shaped National Advisory Committee for Aeronautics (NACA) profiles, depending merely on the variation of Angle of Attack (AOA) such as rudder, hydroplane, and propeller blade. This research is concerned with the deformable hydrofoil that aims at modifying its NACA profile rather than its AOA. However, there is still a lack of knowledge about designing an appropriate deformable hydrofoil. Therefore, a numerical investigation of hydrodynamic characteristics for selected hydrofoils was conducted. After undergoing the 2D numerical analysis (potential flow method) at specific conditions, several NACA profiles were chosen based on the performance of NACA profiles. NACA 0017 was selected as the initial shape for this research before it deformed to the optimized NACA profiles, NACA 6417, 8417, and 9517. The 3D CFD simulations using the finite volume method to obtain hydrodynamic characteristics at 0 deg AOA with a constant flow rate. The mesh sensitivity and convergence study are carried out to get consistent, validated, and reliable results. The final CFD modeled for propeller VP 1304 for open water test numerically. The results found that the performance of symmetry hydrofoil NACA 0017 at maximum AOA is not the highest compared to the other deformed NACA profiles at 0 deg AOA. The numerical open water test showed that the error obtained on K.T., K.Q., and efficiency is less than 8% compared to the experimental results. It shows that the results were in good agreement, and the numerical CFD setting can be used for different deformed profiles in the future.

Influence of a Forward-Facing Step Surface Irregularity on Swept Wing Transition

An experimental study has been carried out to determine the influence of a forward facing-step (FFS) on the laminar–turbulent transition of a swept wing boundary layer. Wind tunnel experiments were conducted at a fixed 3 deg angle of attack and varying Reynolds number between 2.5 and 4.5 million. Moreover, the FFS influence was investigated under unforced (i.e., smooth leading edge) and forced conditions (i.e., using discrete roughness elements). For each test case infrared thermography was used to determine the transition location and spatial organization of the crossflow vortices. In addition, the change in amplification factor was calculated using linear stability theory. Results reveal the importance of considering multiple parameters when estimating the critical FFS height. The unforced cases indicate that one-parameter correlations (i.e., based on the crossflow vortex core height or boundary-layer displacement thickness) might not be sufficient to universally capture the dynamics of these complex flows. Analysis of the forced cases shows that in addition to local parameters (i.e., step height and vortex core height), the FFS influence on transition depends on the stability characteristics of the incoming instability mode. These findings suggest a complex nonlinear interaction between instabilities and surface irregularities, which highlight the need for multiparameter correlations for accurate transition prediction.

Boundary Layer Transition Preflight Experiments in a Mach-6 Quiet Tunnel

Experiments on a subscale model of the U.S. Air Force Office of Scientific Research Boundary Layer Transition flight experiment were made in the Boeing/Air Force Office of Scientific Research Mach-6 Quiet Tunnel. Infrared thermography and temperature-sensitive paint were used to measure the surface temperature and heat transfer and provide images of streamwise streaks with the model at 0, 2, and 4 deg angle of attack. Variations in heating patterns were measured for yaw and roll angles. All smooth body cases remained laminar under quiet flow, but symmetric turbulent wedges were seen in noisy flow. Discrete surface pressure sensors measured possible instabilities on the model surface at locations that correspond to computed instability growth. It was discovered that the geometry was sensitive to forward- and rearward-facing steps near the swept leading edges. Controlled forward- and rearward-facing steps were created and tested. No transition was measured under quiet flow for any rearward-facing step at 0 deg angle of attack. Early transition was measured for the forward-facing steps, and correlations were developed for scaling step effects to the flight vehicle.

Transition and Turbulence Modeling for the Prediction of Cavitating Tip Vortices

Abstract This study evaluates the influence of transition and turbulence modeling on the prediction of wetted and cavitating tip vortices for an elliptical wing, while investigating the numerical errors. Transition modeling increases the quality of numerical predictions since the assumption of a fully turbulent boundary layer, commonly found in literature, contributes to underprediction of the tip vortex cavity size. By applying the local correlation-based transition model (LCTM) and controlling the boundary layer thickness using different turbulent inflow conditions, the pressure in the vortex was found to reduce by 20% for an Angle of Attack (AoA) of 5 deg. The consequent increase in cavity size was found to be of a similar order of magnitude. At 9 deg AoA, transition always occurs just downstream of the leading edge, making this AoA more suitable to investigate the effect of different turbulence modeling approaches. Azimuthal and axial velocity fields are validated against stereographic-particle image velocimetry (S-PIV) measurements. The time-averaged velocity profiles predicted by delayed detached-eddy simulation (DDES) and improved delayed detached-eddy simulation (IDDES) are close to the experiments; however, no velocity fluctuations and vortex dynamics are observed around the vortex. A comparison of wetted and cavitating simulations shows that the cavity leads to a change in the balance between production and destruction of turbulence kinetic energy, which reduces the turbulent diffusion in and around the cavity compared to wetted flow conditions. Consequently, the vapor flow exhibits the characteristics of a potential flow. Whether this is physically plausible remains to be investigated.

Numerical investigations on drag reduction of a civil light helicopter fuselage

The paper presents the numerical investigations for a light civil helicopter fuselage with the purpose of reducing the helicopter drag. Two numerical methods: Reynolds-Averaged Navier-Stokes (RANS) and a hybrid RANS-LES method: improved delayed detached-eddy simulation (IDDES) method are performed for the simulation of a prototype helicopter fuselage. The RANS method is validated by a modular helicopter fuselage with experimental data, while the IDDES method has already been used in other studies and good results have been achieved. The numerical results of the civil helicopter in terms of the aerodynamic coefficients obtained by the two methods are consistent with each other. Compared to the RANS method, the IDDES approach is capable of predicting the large separation flow of the fuselage/components and the small-scale vortices in the flow field. It is found that the hybrid method captures the periodically detached separating vortices behind the skid struts, while the flow at the same position simulated by the RANS method is still attached. Moreover, separating vortices, which develop diagonally upwards, are captured at fuselage/tail-boom junction by the hybrid method, whereas RANS simulation doesn't predict. Subsequently, the design of the helicopter drag reduction is carried out at α=0 deg angle of attack and 0.2 Mach number through shape optimization based on the numerical simulations, and the numerical results are verified by the force tests. It shows that the optimization of the landing skids aims at reducing the strong Karman vortex shedding effect induced by the skids, which greatly decreases the total drag. Combined with the optimization of the fuselage/tail-boom transition, this approach suppresses the separation flow of the fuselage and reduces the pressure drag, and the adjustment of transition slope benefits of up to 1.18% drag reduction, while the extension of transition reduces the drag by 2.64%. The optimization of tail-boom layout demonstrates that lower tail-boom layout can contribute to a drag reduction by 4.37% (RANS) and 3.89% (test), respectively, when compared to the original tail-boom layout.

Parametric Experimental Studies on Supersonic Flow Unsteadiness over a Hemispherical Spiked Body

Experimental studies are carried out to investigate the effects of the geometrical parameters with a drag reducing spike on a hemispherical forebody in a supersonic freestream of at 0 deg angle of attack. The spike length (, 1.0, 1.5, 2.0), spike stem diameter (, 0.12, 0.18), and spike tip shapes are varied, and their influence on the time-averaged and time-resolved flowfield is examined. When increases, a significant reduction in drag is achieved at , whereas the variation in has only a minor effect. The intensity of the shock-related unsteadiness is reduced with an increase in from 0.06 to 0.18, whereas changes in have a negligible effect. The effects of spike tip geometry are studied by replacing the sharp spike tip with a hemispherical one having three different base shapes (vertical base, circular base, and elliptical base). A hemispherical spike tip with a vertical base is performing better by reducing and flow unsteadiness. The dominant spatiotemporal mode arising due to the shock-related unsteadiness is represented through modal analysis of time-resolved shadowgraph images, and the findings are consistent with the other measurements.

Flow Asymmetry in a Y-Shaped Diverterless Supersonic Inlet: A Novel Finding

Extensive wind-tunnel tests were performed on a Y-shaped diverterless supersonic inlet (DSI). All tests were conducted at a free stream Mach number of , the design Mach number for this inlet, and at both zero degrees angle of attack (AOA) and angle of sideslip (AOS). The experiments were performed at various inlet operating conditions comprising supercritical, critical, and subcritical conditions that covered almost all ranges of the engine operating for this DSI. The results showed that the DSI had relatively acceptable performance characteristics when operating at its design condition. A symmetric supersonic flow pattern was observed at both supercritical and critical operating conditions of the present Y-shaped DSI. However, both symmetric and asymmetric supersonic flow patterns were observed during the subcritical operating conditions of the present inlet. At low subcritical operating conditions, the supersonic flow pattern remained symmetric, similar to that of the critical operating condition case. As the inlet experienced higher subcritical operating conditions, the supersonic flow structures became suddenly asymmetric and a new flow pattern was noticed. This flow pattern was seen to be highly unstable; and significant asymmetric shock wave oscillations, boundary-layer thickening, and shock-/boundary-layer interaction were observed for these situations. To the authors’ knowledge, this novel phenomenon of asymmetric flow behavior in a perfectly manufactured symmetric Y-shaped DSI encountering symmetric flow (0 deg AOA and AOS) has not been noticed by previous researchers and, if so, the results have not been published yet.

Effects of Mach Number on the Performance of a Diverterless Supersonic Inlet

Extensive experiments were conducted on a body-integrated diverterless supersonic inlet (DSI). Diverterless supersonic inlets are designed and developed in order to provide both supersonic flow compression and boundary-layer diversion by using a three-dimensional bump in combination with a suitable cowl lip. The present experiments were performed at three different freestream Mach numbers of , 1.65 (the design Mach number), and 1.85, as well as at 0 deg angles of attack and angles of sideslip. To model the performance accurately, the intake was integrated with a typical forebody including a nose with an elliptical cross section. Wind-tunnel tests were conducted at critical, subcritical, and supercritical operating conditions. The results showed that the present DSI has acceptable performance in these operating conditions and is able to provide the required mass flow and static pressure ratios. For all conditions examined in this study, as a significant result, the fixed geometry of the designed DSI showed acceptable performance in the ranges of supersonic Mach numbers tested: ; furthermore, its operation in the subsonic condition of was satisfactory. It should be mentioned that there were no movable parts or an auxiliary flow control system for this intake.