Transonic Mach Numbers Research Articles

ON THE APPLICATION OF BACKGROUND ORIENTED SCHLIEREN TO A TRANSONIC LOW-REYNOLDS TURBINE CASCADE

Abstract Qualitative and quantitative visualizations of transonic turbomachinery flows provide essential information on compressibility effects on boundary layer development, shock-boundary layer interactions and trailing edge flows. Background Oriented Schlieren (BOS) is a relatively new optical technique that allows capturing density gradient fields through turbomachinery cascades and thus, quantitative experimental data to validate transonic and supersonic blade designs. However, only very few experimental works in the open literature have successfully applied BOS to transonic turbine or compressor flows. The current study presents the application of BOS to a transonic low pressure turbine cascade, the VKI SPLEEN C1 cascade for a range of Reynolds numbers (70,000 to 140,000) and transonic Mach numbers (0.90 to 1.00). The linear turbine cascade is tested at the von Karman Institute in the S-1/C high-speed wind tunnel. The test section is instrumented with different BOS optical setups to visualize time-resolved density gradients through the turbine passage and at the trailing edge plane with dedicated field of views. The BOS images are processed using the cross-correlation algorithm, and the optical flow approach, recently introduced in BOS applications to gain spatial resolution and increased sensitivity. Steady-state density gradients of the cascade flow characterize the airfoil boundary layers, wakes and shock waves. The results from the two data reduction methods are assessed and compared against available RANS predictions. Time-resolved measurements reveal the low-frequency motion of weak shock waves generated in the blade passage using POD and spectra analysis.

Benchmark problems for simulating Hyperloop aerodynamics

Hyperloop is proposed as the next generation of sustainable high-speed transport. Recently, an increasing body of literature has been amassed on Hyperloop aerodynamics, however, the vast majority of this work is numerical. Experimentally, there are few relevant studies and none are suitable for validating computational approaches. This paper presents three benchmark cases to provide a framework for computational research and to address this significant gap. Benchmark 1 provides experimental data from existing work on a projectile traveling at Mach 1.1 in ground effect. This incorporates many of the flow characteristics of a Hyperloop system, including (i) transonic Mach numbers, (ii) wall confinement, and (iii) shock formation/reflection. These experimental data are compared to Computational Fluid Dynamics simulations with a very good match seen. Next, Benchmark 2 is proposed which extends these simulations toward a baseline Hyperloop pod design operating in an axisymmetric low-pressure tube environment. This is achieved in stages by adding a full tube, scaling up the domain, reducing the air pressure, and introducing a baseline pod design. It is shown that the enclosed tube environment causes the most significant change in aerodynamic characteristics via flow choking. Nevertheless, a number of aerodynamic similarities remain, compared to Benchmark 1. Finally, Benchmark 3 is proposed to explore the impact of ground clearance of the pod. This aspect has a significant influence on the flow by deflecting the wake and the downstream shock pattern. Furthermore, the drag, downforce, and pitching moment are all found to increase with lower ground clearances.

On the mechanism of transonic buzz passive suppression with global stability analysis

The control surface buzz may occur when the aircraft cruise at transonic Mach numbers, which would may lead to flight accidents. Transonic buzz could be a problem in aircraft design. The triggering of transonic buzz has been investigated in the previous publications. It is essentially a single degree of freedom flutter caused by the coupling of one structural mode and one pre-instability flow mode, thus, type A/B buzz usually occurs close to the buffet onset. With this mechanism, the buzz can be suppressed by improving the flow stability and buffet onset. Three buffet passive control techniques are applied to test the suppression effect on transonic buzz. The relationship between the flow stability and buzz onset is studied to verify the mechanism of buzz. CFD simulation and global stability analysis are applied to compute the flow stability with and without control techniques. Results show that, three control techniques can improve the flow stability and postpone the buffet onset. Meanwhile the buzz can be effectively suppressed. With the aeroelastic ROM and the root loci method, the coupling process between the structural mode and the controlled flow modes are revealed, indicating the buzz suppression effects are positively correlated with its improvement of the flow stability with control. This study can help to understand the mechanism of buzz deeply and propose the new thought of buzz suppression.

Alleviation of SWBLI over the payload of a launch vehicle by change of nose shape

Wind tunnel tests on a typical heavy-lift launch vehicle configuration featuring a core vehicle flanked by two strap-on boosters showed very high levels of unsteady pressures over the payload region at transonic Mach numbers when the nose had a cone-cylinder shape. This was due to very high levels of Shock Wave Boundary Layer Interactions (SWBLI) associated with the conical shape, which caused alternating flow featuring the formation of a unsteady λ-shock system over the payload region and a pair of counter-rotating vortices. The phenomenon was found to strongly depend on the proximity of the noses of the strap-on boosters to the boat tail downstream of the payload region. When the conical nose was replaced by an ogive-shaped nose, significant reduction in SWBLI occurred, with dramatic reduction of shock oscillations, pressure fluctuations, replacement of alternating flow by steady flow and disappearance of the vortex-pair. Time-averaged pressure fluctuations indicate reduction in pressure gradients due to a more gradual expansion of flow over the ogive nose compared to conical nose, an observation also supported by CFD simulations. The alleviation of SWBLI with ogive nose was so effective that the presence or absence of strap-on boosters had practically no effect on pressure fluctuations over the payload region. Unlike the conical shape, the ogive nose cone shape seems to meet both the shape and pressure parameter-criteria of NASA to ensure buffet-free environment.

RANS Prediction of Losses and Transition Onset in a High-Speed Low-Pressure Turbine Cascade

Current trends in aero-engine design are oriented at designing high-lift low-pressure turbine blades to reduce engine weight and dimensions. Therefore, the validation of numerical methods able to correctly capture the boundary layer transition at cruise conditions with a steady inflow for high-speed blades is of great relevance for turbine designers. The present paper details numerical simulations of a novel open-access high-speed low-pressure turbine test case that are performed using RANS-based transition models. The test case is the SPLEEN C1 cascade, tested in transonic conditions at the von Karman Institute for Fluid Dynamics. Both physics-based and correlation-based transition models are employed to predict blade loading, boundary layer characteristics, and wake development. 2D simulations are run for a wide range of operating conditions ranging from low to high transonic Mach numbers (0.7–0.95) and from low to moderate Reynolds numbers (70,000–120,000). The γ-Re˜θt transition model shows a good performance over the whole range of simulated operating conditions, thus demonstrating a good capability in both reproducing blade loading and average losses, although the wake’s width is underestimated. This leads to an overestimation of the total pressure deficit in the center of the wake which can exceed experimental measurements by more than 50%. On the other hand, the k-ν2-ω model achieves satisfactory results at Ma6,is = 0.95, where the boundary layer state is affected by the presence of a weak shock impinging on the blade suction side which thickens the boundary layer, leading to a predicted shape factor equal to five, downstream of the shock. However, at low and moderate Mach numbers, the k-ν2-ω model predicts long or open separation bubbles contrary to the experimental findings, thus indicating insufficient turbulence production downstream of the boundary layer separation. The slow boundary layer transition in the aft region of the suction side that is exhibited by the k-ν2-ω model also affects the prediction of the outlet flow, featuring large peaks of a total pressure deficit if compared to both the experimental measurements and the γ-Re˜θt predictions. For the k-ν2-ω model, the maximum overestimation of the total pressure deficit is approximately 60%.

Effect of Flare Angle and Natural Ventilation on The Aerodynamic Characteristics of a Typical Re-Entry Body At Subsonic and Transonic Mach Numbers

Aerodynamic force and moment measurements were carried out on models of a typical re-entry body, featuring a blunt nose, conical body and flare, to assess the effect of flare angle on the overall aerodynamic characteristics. In one of the configurations, the junction between the cone and the flare was vented out to the base of the model to assess the effect of ventilation on the aerodynamic characteristics. It was expected that the stability characteristics would improve, by virtue of improved contribution from the flare to the normal force of the configuration. However, measurements did not show the expected trends. Instead, there was a remarkable reduction in the axial force with ventilation, by as much as 25% all across the Mach number and angle of attack range. Base pressure measurements indicated that in the low Mach number range pressure rise occured. However, in the high transonic Mach number range, even though the force measurements showed drag reduction, there was no significant rise in base pressure. Ventilation showed delay in transonic drag rise. It is therefore suggested that the overall pressure field is modified by ventilation, which results in drag reduction of the configuration.

Numerical Flowfield Visualization Over Heat Shield Of Satellite Launch Vehicle at Mach 0.8 - 3.0

The paper presents a numerical flowfield visualization over a typical payload shroud of a satellite launch vehicle at Mach number range of 0.80 -3.0 and Reynolds number range of 33 x 106 - 46 x 106 /m. The numerical simulation over the bulbous heat shield is carried out by solving time-dependent compressible axisymmetric turbulent Reynolds-averaged NavierStokes equations. The closure of these equations is obtained employing the Baldwin-Lomax turbulence model. A three-stage Runge-Kutta time-stepping scheme has been used in conjunction with finite-volume discretization of the computational domain. The flowfield features over the bulbous heat shield have been analyzed from the velocity vector and the density contour plots. Flow separation on the payload shroud due to the normal shock wave is observed at Mach number 0.80 and 0.90. The normal shock movement on the heat shield is simulated for various transonic Mach number. A recirculation zone of flowfield is formed in the boat-tail region of the heat shield. The vorticity formation ahead of the heat shield moves close to the heat shield with the increasing transonic Mach number, which is observed in the density contours plots of the flowfield. Shock stand-off distance at the supersonic speed is calculated and compared with asymptotic formula of Frank and Zierep. They are found in good agreement.

A Reconstructive Approach for Design of a Planar Supercritical Wing

Effort is made to design an efficient planar wing that can support transonic pressure distribution. Wing for transport airplane role is considered. First optimal warp is generated for minimum induced drag for a Mach number that is just below transonic Mach number. Optimal warp thus generated is split into twist and camber. The resulting camber is used to form a planar wing. The camber of this aerofoil is varied along span. Twist is no more considered. Aerofoil thickness is superimposed on this camber. Mach number is increased till supercritical flow appears.The whole process is reconstructive in approach. Results obtained are very encouraging and are application oriented. Pattern of supercritical pressure distribution is studied from point of view of suction peaks, shock wave location, and aft loading intensity occurring towards trailing edge. Main effort of our exercise is to create shock free transonic wing.

Experimental Investigation for Subsonic, Transonic, and Supersonic Performances of Double-Swept Waverider

The wind-tunnel experiments for the double-swept waverider were carried out in subsonic, transonic, and supersonic states, aiming to validate its potential advantages in wide-speed performances. The test configurations were generated using the planform-customized waverider design, which was developed from the osculating-cone method. Two experimental models of the double-swept waveriders with bend and cusp heads were fabricated, and a single-swept configuration was built as a comparable model. A series of experiments were performed in the FD-06 wind tunnel of the China Academy of Aerospace Aerodynamics at a subsonic Mach number of ; transonic Mach numbers of , 0.8, and 0.9; and a supersonic Mach number of . The results from the wind-tunnel tests were examined by combining the numerical simulations. As the Mach number increased, the lift-to-drag () ratios of the double-swept waveriders decreased slightly when ; but when , their dropped sharply. Moreover, the double-swept waveriders were unstable in the longitudinal direction in the subsonic state, and their longitudinal stability was obviously improved as the Mach number went up. When the Mach number was greater than 0.9, they became statically stable. Compared with the single configuration, the double-swept waverider exhibited increased in subsonic and transonic states, as well as improved longitudinal stability in the wide-speed range, but it exhibited a decreased in the supersonic state.

Shock Buffet and Associated Fluid–Structure Interactions of the Benchmark Supercritical Wing

This study focuses on the fluid–structure interactions of the Benchmark Supercritical Wing (BSCW) at buffet conditions (transonic flows and moderate angles of attack). It presents unsteady Reynolds-averaged Navier–Stokes simulations of the fixed BSCW, the BSCW undergoing prescribed pitch motion, and the BSCW suspended on springs in one and two degrees of freedom. The study’s objectives are to 1) establish the BSCW buffet properties and 2) investigate the aeroelastic and prescribed-motion responses of the wing at buffet conditions. The results indicate that the BSCW buffet is predominantly two-dimensional in terms of its characteristics, although it is greatly affected by the tip flow. The response to prescribed pitch motion and the aeroelastic response of the spring-suspended wing display similar phenomena, where the response in the buffet frequency is suppressed once the oscillation amplitude exceeds some threshold, and the structural and aerodynamic responses are solely at the structural frequency. At the studied conditions, the BSCW undergoes stall flutter, dominated by pitch oscillations, compared to a bending-torsion coalescence flutter at lower angles of attack. We conclude that the shock buffet does not drive the BSCW flutter instability. However, both the buffet and stall flutter develop at similar flow conditions of detached boundary layer behind a strong shock at transonic Mach numbers and moderate angles of attack.

Experimental Investigation of Transonic Shock Buffet on an OAT15A Profile

Self-sustained shock wave oscillations on airfoils, commonly defined as shock buffet, can occur under certain combinations of transonic Mach numbers and angles of attack (AoAs) due to the interaction between the shock and the separated boundary layer. To improve the understanding of this complex phenomenon, the flow over a supercritical profile (OAT15A) was experimentally investigated for a fixed Reynolds number of and numerous aerodynamic conditions within the ranges of and . Deformation and force measurements were used to assess the actual rigidity of the model and its interaction with the flow. The tracking of the shock location by means of background-oriented schlieren allowed for studying the shock features and the frequency content of the buffet flows. Furthermore, the inversion of shock motion and buffet onset, which are referred to as buffet boundaries, were estimated. The inversion of shock motion proved to be a necessary but not sufficient condition for buffet onset. The trends of the results showed a good agreement with the relevant literature cases. However, the buffet amplitude was smaller, and buffet onset occurred at considerably higher AoAs. The comparison of the literature results also revealed a general sensitivity of buffet features to both numerical and experimental parameters. For this reason, the influence of the boundary-layer suction at the vertical walls and the gap flow at the side windows on the buffet features was examined. The buffet frequencies and amplitudes were slightly affected, but the buffet boundaries appeared to be virtually insensitive to these factors. Given the large number of investigated aerodynamic conditions, these results are valuable for validation purposes of computational fluid dynamics simulations.

Influence of Fluid Viscosity and Compressibility on Nonlinearities in Generalized Aerodynamic Forces for T-Tail Flutter

The numerical assessment of T-tail flutter requires a nonlinear description of the structural deformations when the unsteady aerodynamic forces comprise terms from lifting surface roll motion. For linear flutter, a linear deformation description of the vertical tail plane (VTP) out-of-plane bending results in a spurious stiffening proportional to the steady lift forces, which is corrected by incorporating second-order deformation terms in the equations of motion. While the effect of these nonlinear deformation components on the stiffness of the VTP out-of-plane bending mode shape is known from the literature, their impact on the aerodynamic coupling terms involved in T-tail flutter has not been studied so far, especially regarding amplitude-dependent characteristics. This term affects numerical results targeting common flutter analysis, as well as the study of amplitude-dependent dynamic aeroelastic stability phenomena, e.g., Limit Cycle Oscillations (LCOs). As LCOs might occur below the linear flutter boundary, fundamental knowledge about the structural and aerodynamic nonlinearities occurring in the dynamical system is essential. This paper gives an insight into the aerodynamic nonlinearities for representative structural deformations usually encountered in T-tail flutter mechanisms using a CFD approach in the time domain. It further outlines the impact of geometrically nonlinear deformations on the aerodynamic nonlinearities. For this, the horizontal tail plane (HTP) is considered in isolated form to exclude aerodynamic interference effects from the studies and subjected to rigid body roll and yaw motion as an approximation to the structural mode shapes. The complexity of the aerodynamics is increased successively from subsonic inviscid flow to transonic viscous flow. At a subsonic Mach number, a distinct aerodynamic nonlinearity in stiffness and damping in the aerodynamic coupling term HTP roll on yaw is shown. Geometric nonlinearities result in an almost entire cancellation of the stiffness nonlinearity and an increase in damping nonlinearity. The viscous forces result in a stiffness offset with respect to the inviscid results, but do not alter the observed nonlinearities, as well as the impact of geometric nonlinearities. At a transonic Mach number, the aerodynamic stiffness nonlinearity is amplified further and the damping nonlinearity is reduced considerably. Here, the geometrically nonlinear motion description reduces the aerodynamic stiffness nonlinearity as well, but does not cancel it.

Computational Fluid Dynamics Analysis and Design of Payload Shroud of Satellite Launch Vehicle

The main focus of the present paper is to computational fluid dynamics analysis and design of payload fairing of satellite launch vehicle at freestream Mach number range of 0.6 - 3.0. Initially, time-dependent compressible three-dimensional Euler equations are solved employing a finite volume discretization method with a multi-stage Runge-Kutta time-stepping scheme to compute surface pressure and aerodynamic coefficients at various payload fairing and at angle of attack up to 5o with an increment of 1o. Payload fairing dimensions are selected that satisfies permissible structure load on satellite launch vehicle. Detailed flowfield simulation is carried out on the selected payload fairing employing axisymmetric compressible Reynolds-average Navier-Stokes equations to assess unsteady flowfield characteristics. The numerical simulations are used to locate terminal shock on the payload fairing at transonic Mach number. Unsteady flow characteristics are used to compute acoustic load. Shock standoff distances at supersonic speeds are tabulated and compared with the analytical solution. Schlieren images and oil flow pictures are compared with experimental results and in good agreement. Aerodynamic shape optimization of satellite launch vehicle payload fairing shape has been performed to satisfy structural load at maximum drag and dynamic pressure.

Effect of Wing Geometric Parameters on Flutter Speed

Fluid-Structure Interaction (FSI) of an aircraft wing is given much importance in recent days due to safety in operation while reducing structural weight of the wing. Any design modification or newer design of an aircraft is expected to have improvements in overall performance. The geometric variation of the wing for aerodynamic performance is given importance before considering it for any major changes. It is important to know the effect of wing modification on flutter speed. Analytical and semi-empirical tools were developed in earlier days to estimate the flutter speed. Currently, coupled Computational Fluid Dynamics (CFD) and Computational Structural Dynamics (CSD) simulations are possible due to development in numerical methods and availability of increased computing power. Ansys-Multi Field Solver is used in this study. A well-known AGARD 445.6 wing geometry is considered for transonic Mach number of 0.901. The predicted onset flutter speed, 297m/s, has shown good agreement with the experimental data. A parametric study on reduction of 10% wing thickness predicted 1% reduction in flutter speed and increase of 10% wing thickness predicted, 5% more on flutter speed. Increase in wingspan by 10% has predicted reduction on the flutter speed of 2% and decrease in wingspan by 10% has predicted increased flutter speed by 5%. Increase in wing sweep by 10% increases flutter speed by 2% and reduction in sweep by 10% has decreased flutter speed by 5%.

Analysis of Averaging Methods for Nonuniform Total Pressure Fields

Abstract A common objective in the analysis of turbomachinery components (nozzle guide vanes (NGVs) or rotor blades, for example) is to calculate performance parameters, such as total pressure or kinetic energy (KE) loss coefficients, from measurements in a nonuniform flow-field. These performance parameters can be represented in a range of ways. For example, line-averages used to compare performance between different radial sections of a 3D component; plane-averages used to assess flow (perhaps loss coefficient) development between different axial planes; and fully mixed-out values used to determine the total loss associated with a component. In the literature, the weighting method used for line- and plane-averaging (e.g., area, volume flow, mass flow, or entropy-flux) is sometimes regarded as an unimportant issue. Indeed, many authors neglect to even state which weighting method was used in their work. In certain low-speed test cases, or where measurements are made a long distance from the component, the nonuniformity in the flow will be relatively small and the practical difference between different weighting methods may be negligible. However, in high-speed applications or for measurements close to a component trailing edge, this becomes increasingly unlikely. In this paper, we compare a range of methods for calculating aerodynamic performance parameters—for example, the kinetic energy loss coefficient—including plane-average methods with different weighting schemes and several mixed-out methods. We analyze the sensitivities of the different methods to the axial location of the measurement plane, the radial averaging range, and the exit Mach number. We use high-fidelity experimental data taken in several axial planes downstream of a cascade of engine parts: high-pressure (HP) turbine NGVs operating at transonic Mach number. The experimental data are complemented by computational fluid dynamics (CFD). We discuss the underlying physical mechanisms which give rise to the observed sensitivities. The objective is to provide guidance on the accuracy of each method in a relevant, practical application.

Computational fluid dynamics for turbomachinery technologies with a focus on high-speed rotor blades

NASA's SR3 8-bladed rotor with a rotational speed of 6350 rpm, an advance ratio of 3.6, and a transonic Mach Number of 0.8 is utilized to evaluate the aerodynamic and rotor effectiveness predictive ability of a CFD simulation using the Ansys commercial software. The experimentally tested model [1] included eight 45° swept blades that were tested at a design operating altitude of 10.68 km (35,000 ft.) and it is the one with the most data in the public domain. In the model, blade sweep and the distinct contour of the spinner and nacelle design were implemented to reduce compressibility losses. The primary goal of this research is to compare performance predictions with the experimental results for power coefficient and efficiency evaluation. Analyses are carried out on a single rotating zone separated by a sliding mesh boundary. We also look at the accuracy of recording downward velocities and swirling angles in the propeller's wake, with the goal to use the RANS technique to investigate how these wake parameters react with aerodynamic surface and the results are compared with the experimental results. Initial computational findings demonstrate increased correlation with wind tunnel measurements at 63.3° degrees pitch.

Control of shock-induced vortex breakdown on a delta-wing-body configuration in the transonic regime

AbstractShock-induced vortex breakdown, which occurs on the delta wings at transonic speed, causes a sudden and significant change in the aerodynamic coefficients at a moderate angle-of-attack. Wind-tunnel tests show a sudden jump in the aerodynamic coefficients such as lift force, pitching moment and centre of pressure which affect the longitudinal stability and controllability of the vehicle. A pneumatic jet operated at sonic condition blown spanwise and along the vortex core over a 60° swept delta-wing-body configuration is found to be effective in postponing this phenomenon by energising the vortical structure, pushing the vortex breakdown location downstream. The study reports that a modest level of spanwise blowing enhances the lift by about 6 to 9% and lift-to-drag ratio by about 4 to 9%, depending on the free-stream transonic Mach number, and extends the usable angle-of-attack range by 2°. The blowing is found to reduce the magnitude of unsteady pressure fluctuations by 8% to 20% in the aft portion of the wing, depending upon the method of blowing. Detailed investigations carried out on the location of blowing reveal that the blowing close to the apex of the wing maximises the benefits.

Effect of Aerospike on Unsteady Transonic Flow over a Blunt Body

Blunt-nosed launch vehicles featuring large nose-cone angles experience high levels of pressure fluctuations over the payload region at transonic Mach numbers due to shock-wave/boundary-layer interactions. The main cause for this phenomenon appears to be flow instability associated with separated flow and formation of a vortex pair. Interactions between the induced velocity of the vortex pair and oncoming mean flow cause oscillations of the -shock system and high levels of fluctuating pressures. An aerospike causes flow separation at the nose and reattachment of the shear layer downstream, energizing the boundary layer. Consequently, flow separation and vortex formation are prevented: shock oscillations are stabilized. Dramatic reductions in the pressure fluctuations, around 95–35% in the low-frequency range, are observed along the payload region at small angles of attack. The observations are based on wind-tunnel tests involving unsteady pressure measurements, surface-flow patterns, and high-speed shadowgraph recordings on a blunt nose cone with various cone angles.

Parametric Study of Separation and Reattachment in Transonic Airfoil Flows

Shock and boundary-layer interactions in transonic airfoils are investigated by means of wall-modeled large-eddy simulation. A distinctive flow model, organized and chaotic, occurs in this regime driven by the transition from laminar to turbulent flow, and therefore the appearance of coherent structures such as the Kelvin–Helmholtz instability. This phenomenon is crucial for designing transonic airfoils because it leads to a high rise in drag. Experimental observations show that the main variables of transonic airfoil flow are the freestream Mach number, the Reynolds number, and boundary-layer thickness. Simulations are performed to investigate the relation of various components of the shock/boundary-layer interaction, that is, the pressure rise to shock-induced separation, the length of the separation bubble, and the rear separation to the unit Reynolds number and the boundary-layer condition upstream of the shock. Flow around NACA0012 is investigated in a range of transonic Mach numbers at various unit Reynolds numbers. It is found that the bubble length on the airfoil changes rapidly with relatively small changes in the freestream Mach number. Simulation results show that the separation length decreases with increasing of the unit Reynolds number; however, increasing the boundary-layer thickness results in an increase in the bubble’s extent. Evaluation of the rear-separation region shows that, as the Reynolds number increases, the length of the rear separation declines at a given Mach number.

Vortex-Generating Shock Control Bumps for Robust Drag Reduction at Transonic Speeds

The 3D contour bump has been integrated with vane-type vortex generators to weaken the shock wave and suppress potential flow separation at higher transonic speeds. The integrated vortex-generating shock control bump is parameterized as a whole, in a manner that the contour bump is designed for shock control and the vortex-generating fins are integrated to the contour bump for separation control. This allows the balance between the two factors affecting the total wing drag. The optimization studies based on the Reynolds-averaged Navier–Stokes equations have been carried out to find the optimal vortex-generating bump designs at the given design conditions. Single-point and multipoint global optimizations are carried out to search the optimum parameters at the corresponding design points. It is found that the vortex-generating bump can further reduce the total drag at the higher transonic Mach numbers due to the alleviation of after-bump streamwise separation by a pair of counter-rotating vortices generated. The inclusion of the vortex generators in the bump design allows for a better balance of the design in controlling both the shock strength and the flow separation after the bump at higher Mach numbers. A multipoint optimization leads to a robust vortex-generating bump design for a range of Mach numbers. The distinctive vortical flow structures induced by the vortex-generating bump are highlighted and discussed.