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The results of the swept wing boundary layer stability investigation are presented for the case, when the wing surface has a region of gas suction through the wall normal to the surface, while the wing is in Mach number 2 flow. In the flow regime considered the predominant boundary layer instability type is the crossflow instability. The gas suction effect on the development of unstable modes in the boundary layer is investigated using the linear stability theory and direct numerical modeling. The numerical modeling of laminar (undisturbed) flow fields with regions of gas suction and disturbed flow fields is carried out by integrating Navier–Stokes equations. An analysis within the framework of the linear stability theory is performed using the -method. The suction region location is varied with conservation of the integral intensity. It is shown that the mode instability growth can be considerably suppressed at the expense of an optimal disposition of the suction region. [ABSTRACT FROM AUTHOR]

In this work, the crossflow instability modes on a 7° half-angle cone with 6° angle of attack at Mach 6 have been experimentally studied. To begin with, the complete laminar-to-turbulence transition process is acquired by infrared thermography, and the heat flux streaks caused by the stationary crossflow vortices are distinct over the areas of leeward side. Further wavelet analysis reveals a dominant wavenumber of the stationary crossflow vortices, which is centered at about k = 50. Subsequently, the spatial distribution characteristics of the traveling crossflow instability (f = 20–50 kHz) and high-frequency instabilities (f = 100–500 kHz) are acquired between the azimuthal angle Φ = 90°–150° (windside to leeside) through wall mounted pressure sensors. The results of spectral and wavelet analysis indicate that the two types of instability are the traveling crossflow instability and its secondary instability, respectively. Further bispectral analysis is used to show that these high-frequency waves undergo several quadratic phase-coupled interactions with themselves to produce harmonics, as well interact with low-frequency waves that results in spectral broadening. • The wavenumber of stationary crossflow waves in noisy condition is similar to that in the quiet flow condition. • The secondary instabilities of traveling crossflow waves are distinct before transition. • The nonlinear interactions between high-frequency and low-frequency instabilities are important transition mechanisms. [ABSTRACT FROM AUTHOR]

Surface suction provides an efficient way to delay boundary layer transitions. In order to verify the suction effects and determine the mechanism of suction control in transonic swept wing boundary layers, wind tunnel transition measurements in a hybrid laminar flow control (HLFC) wind tunnel model uses an infrared thermography technique in the Aircraft Research Association (ARA) 2.74 m × 2.44 m low turbulence level transonic wind tunnel. Based on the experimental data of stationary crossflow dominant transitions without and with surface suction in transonic swept wing boundary layers, in this paper, the effects on the receptivity and linear and nonlinear evolution of stationary crossflow vortices have been analyzed with the consideration of curvature. Theoretical analysis agreed with the experimental observations in regard to the transition delay caused by boundary layer suction near the leading-edge region. [ABSTRACT FROM AUTHOR]

Theoretical assessments of the crossflow (CF) stabilization due to flow slip produced by small grooves on a swept supersonic wing are performed using the linear theory for inviscid flow, the local similar approximation of the boundary layer flow, the slip boundary conditions on the grooved surface and the linear stability theory. The e N computations for stationary CF mode predict that spanwise-invariant grooves with their half-period equal to 0.25 of the boundary-layer displacement thickness can delay the CF-induced transition onset by about 10% on a 30 ∘ swept wing having a parabolic airfoil of 5% thickness ratio, at freestream Mach number 2. It is concluded that the groove laminarization concept deserves further studies. [ABSTRACT FROM AUTHOR]

Laminar-to-turbulent transition in a swept flat-plate boundary layer is caused by the breakdown of the crossflow vortex via high-frequency secondary instability and is promoted by the wall-surface roughness and the freestream turbulence (FST). Although the FST is characterized by its intensity and wavelength, it is not clear how the wavelength affects turbulent transitions and interacts with the roughness-induced transition. The wavelength of the FST depends on the wind tunnel or in-flight conditions, and its arbitrary control is practically difficult in experiments. By means of direct numerical simulation, we performed a parametric study on the interaction between the roughness-induced disturbance and FST in the Falkner–Skan–Cooke boundary layer. One of our aims is to determine the critical roughness height and its dependence on the turbulent intensity and peak wavelength of FST. We found a suppression and promotion in the transition process as a result of the interaction. In particular, the immediate transition behind the roughness was delayed by the long-wavelength FST, where the presence of FST suppressed the high-frequency disturbance emanating from the roughness edge. Even below the criticality, the short-wavelength FST promoted a secondary instability in the form of the hairpin vortex and triggered an early transition before the crossflow-vortex breakdown with the finger vortex. Thresholds for the FST wavelengths that promote or suppress the early transition were also discussed to provide a practically important indicator in the prediction and control of turbulent transitions due to FST and/or roughness on the swept wing. [ABSTRACT FROM AUTHOR]

Surface suction provides an efficient way to delay boundary layer transitions. In order to verify the suction effects and determine the mechanism of suction control in transonic swept wing boundary layers, wind tunnel transition measurements in a hybrid laminar flow control (HLFC) wind tunnel model uses an infrared thermography technique in the Aircraft Research Association (ARA) 2.74 m × 2.44 m low turbulence level transonic wind tunnel. Based on the experimental data of stationary crossflow dominant transitions without and with surface suction in transonic swept wing boundary layers, in this paper, the effects on the receptivity and linear and nonlinear evolution of stationary crossflow vortices have been analyzed with the consideration of curvature. Theoretical analysis agreed with the experimental observations in regard to the transition delay caused by boundary layer suction near the leading-edge region.

Boundary-layer transition on swept wings is dominantly caused by the crossflow instability, which is expected to be suppressed by placing artificial roughness elements near the leading edge. It is however difficult to find the optimal roughness shape by using direct numerical simulation (DNS), because a lot of computations are required for assessing a suppression effect due to one roughness shape. In this study, we develop an efficient method to evaluate the suppression effect for a series of roughness shapes by changing a shape parameter quasi-statically and observing the subsequent change of the crossflow mode at a downstream position. Since the mode grows spatially as convective instability, we need to allow for the delay time for the change in the shape to cause the change in the mode. This method is demonstrated for optimizing the height and angle of sinusoidal roughness elements. By applying a volume penalization (VP) method, the height and angle are changed slowly in DNS, where the initial values, rates of change and permeability of the VP method should be chosen appropriately to reproduce the correct results for the fixed shapes. The method developed here shows that the suppression (or laminarizing) effect tends to be improved as the height is increased, but there is a critical height at which flow tripping occurs. Both the laminarization effect and the critical height vary greatly depending on the angle. This result suggests the optimal roughness shape, considering the effectiveness and robustness. For laminar flow control, this method is useful for optimizing the wall roughness shape.

Hypersonic boundary-layer flows over a circular cone at a moderate yaw angle can support strong crossflow instability away from the windward and leeward rays on the plane of symmetry. Due to the more efficient excitation of stationary crossflow vortices by surface roughness, a possible path to transition in such flows corresponds to rapid amplification of the high-frequency instabilities sustained in the presence of finite amplitude stationary crossflow vortices. This paper presents a computational analysis of crossflow instability over a 7-degree half-angle, yawed circular cone in a Mach 6 free stream. Specifically, the nonlinear evolution of an azimuthally localized crossflow vortex pattern and the linear amplification characteristics of high-frequency instabilities evolving in the presence of that pattern are described for the first time. Focusing on the azimuthally compact vortex pattern allows us to overcome significant limitations of the prior secondary instability analyses of azimuthally inhomogeneous boundary layer flows. A comparison between plane-marching parabolized stability equations and direct numerical simulations (DNS) reveals favorable agreement in regard to mode shapes, most amplified disturbance frequencies, and the N-factor evolution. In contrast, the quasiparallel predictions are found to result in a severe underprediction of the N-factors. The most amplified high-frequency instabilities are found to originate from Mack's second mode waves sustained within the upstream region of nearly unperturbed, quasi-homogeneous boundary layer. [ABSTRACT FROM AUTHOR]

Boundary layer transition is an important factor in the design process of hypersonic vehicles, and crossflow instability is a critical mode. In this paper, the traveling crossflow instability in a laminar boundary layer at Mach 6 is experimentally investigated. Experiments were performed over a 7 ∘ half-angle cone at 6 ∘ angle of attack, with the unit Reynolds number of R e ∞ = 5. 75 × 1 0 6 m − 1. Flush-mounted fast-response pressure transducers and time-averaging temperature-sensitive paint were used to capture the traveling crossflow instability and the stationary crossflow instability, respectively. The traveling crossflow waves with characteristic frequencies between 10 and 25 kHz are detected in the planes with the azimuthal angle between 23 ∘ and 143 ∘ away from the leeward ray. The three-dimensional amplitude growth was obtained. In general, the amplitude growth of the traveling crossflow instability is roughly leeside-forward and windside-aft. Although the traveling crossflow instability grows earlier on the leeward side, it has a larger amplitude on the windward side. Compared with the stationary mode, the traveling mode grows faster and reaches the highest amplitude earlier. In addition, a high-frequency instability with characteristic frequencies between 100 and 200 kHz is observed. It may be the secondary instability of the traveling crossflow waves, but whether it is true or not needs further experimental and numerical studies. • In a hypersonic flow, the three-dimensional growth of the traveling crossflow instability on a yawed cone was measured. • Compared with the stationary crossflow instability, the traveling crossflow instability grows faster and reaches the highest amplitude earlier. • A high-frequency instability with f = 100–200 kHz was found, and it may be the secondary instability of the traveling crossflow waves. [ABSTRACT FROM AUTHOR]

Laminar-to-turbulent transition in a swept flat-plate boundary layer is caused by the breakdown of the crossflow vortex via high-frequency secondary instability and is promoted by the wall-surface roughness and the freestream turbulence (FST). Although the FST is characterized by its intensity and wavelength, it is not clear how the wavelength affects turbulent transitions and interacts with the roughness-induced transition. The wavelength of the FST depends on the wind tunnel or in-flight conditions, and its arbitrary control is practically difficult in experiments. By means of direct numerical simulation, we performed a parametric study on the interaction between the roughness-induced disturbance and FST in the Falkner–Skan–Cooke boundary layer. One of our aims is to determine the critical roughness height and its dependence on the turbulent intensity and peak wavelength of FST. We found a suppression and promotion in the transition process as a result of the interaction. In particular, the immediate transition behind the roughness was delayed by the long-wavelength FST, where the presence of FST suppressed the high-frequency disturbance emanating from the roughness edge. Even below the criticality, the short-wavelength FST promoted a secondary instability in the form of the hairpin vortex and triggered an early transition before the crossflow-vortex breakdown with the finger vortex. Thresholds for the FST wavelengths that promote or suppress the early transition were also discussed to provide a practically important indicator in the prediction and control of turbulent transitions due to FST and/or roughness on the swept wing.

An experimental study on the traveling crossflow instability over a 60 ∘ swept flat plate was conducted. The Mach number is 6, the angle of attack of the model is 5 ∘ . The traveling crossflow waves and the secondary instability of the traveling crossflow waves were visualized by nano-tracer-based planar laser scattering (NPLS) technique. In the spanwise NPLS images, the traveling crossflow waves appeared as regular strikes, and the secondary instability appeared as small eddies attached to strikes. The wavelet transform was used to study the wavelength contents of the traveling crossflow waves. The most amplified wavelength is stable before the secondary instability happening, which is around 12 mm at R e ∞ = 3.45 × 10 6 m - 1 . Besides, the Reynolds number effects on the boundary layer transition and traveling crossflow instability were discussed. [ABSTRACT FROM AUTHOR]

Based on the database from linear stability theory (LST) analysis, a local amplification factor transport equation for stationary crossflow (CF) waves in low-speed boundary layers was developed in 2019. In this paper, the authors try to extend this transport equation to compressible boundary layers based on local flow variables. The similarity equations for compressible boundary layers are introduced to build the function relations between non-local variables and local flow parameters. Then, compressibility corrections are taken into account to modify the source term of the transport equation. Through verifications of different sweep angles, Reynolds numbers, angles of attack, Mach numbers, and different cross-section geometric shapes, the rationality and correctness of the new transport equation established in this paper are illustrated. [ABSTRACT FROM AUTHOR]

The paper presents results of experimental study for a flow on the windward side of a swept wing with disturbance generators installed on the surface. These generators are 3D roughness elements with the height comparable to the boundary layer thickness. The method of liquid crystal thermography was used for studying the impact of roughness elements with different heights on the boundary layer. There exists a zone of maximal susceptibility of the flow to the disturbance generated past the roughness element on the wing surface. [ABSTRACT FROM AUTHOR]

This paper describes work carried out within the European Union (EU)-Russia Buterfli project to look at the control of transition-causing "target" stationary cross flow vortices, by the use of distributed plasma actuation to generate sub-dominant "killer" modes. The objective is to use the "killer" modes to control the "target" modes through a non-linear stabilizing mechanism. The numerical modelling and results are compared to experimental studies performed at the TsAGI T124 tunnel for a swept plate subject to a favorable pressure gradient flow. A mathematical model for the actuator developed at TsAGI was implemented in a linearized Navier–Stokes (LNS) solver and used to model and hence predict "killer" mode amplitudes at a measurement plane in the experiment. The LNS analysis shows good agreement with experiment, and the results are used as input for non-linear parabolized stability equation (PSE) analysis to predict the effect of these modes on crossflow transition. Whilst the numerical model indicates a delay in transition, experimental results indicated an advance in transition rather than delay. This was determined to be due to actuator-induced unsteadiness arising in the experiment, resulting in the generation of travelling crossflow disturbances which tended to obscure and thus dominate the plasma stabilized stationary disturbances. [ABSTRACT FROM AUTHOR]

Highlights • Assessment of three tansition models for simulations of HIAD is conducted. • Crossflow effects on performances of transition models for HIAD are analyzed. • Crossflow instability mode and the first disturbance mode dominate the transition. Abstract Hypersonic Inflatable Aerodynamic Decelerator (HIAD) has shown its great potential for future planetary explorations. However, the HIAD surface deflections could both promote boundary-layer transition early and augment heating levels sharply, which poses challenges for survivability of a Thermal Protection System (TPS). The goal of this work is to assess different transition models for the prediction of hypersonic transitional flows over scalloping deformed HIAD surface and to seek the critical factor for their capabilities to provide a reference for the application and advancement. Three representative transition models are considered: γ-Re θ , k-ω-γ and k T -k L -ω. The results show that the undulating surface causes flow separations and reattachments in valleys and strong crossflow along the leeward. k-ω-γ model can correctly predict both the shapes and locations of the transition onsets. The transition zone predicted by γ-Re θ model is much small, while k T -k L -ω is incapable of transition prediction for such a complex and irregular configuration. Moreover, this study reveals that the crossflow instability plays a dominant role in the transition. The crossflow Reynolds number Re cf , whose distributions are approximately consistent with the transition zone, could be a feasible crossflow strength indicator for the transition onsets on the undulating surface. Once k-ω-γ model excludes effects of crossflow instability, it predicts incorrect transition results for the deformed surface of HIAD. Besides, a detailed analysis shows that both the crossflow instability mode and the first disturbance mode in valleys are the major contributors to the transition on the leeward surface. Near the leeward ray, the transition is mainly determined by the first disturbance mode. [ABSTRACT FROM AUTHOR]

The possibility of laminarization of flow around a swept wing using surface microreliefs in the form of riblets and separate plates (winglets) set at an angle to external streamlines is theoretically investigated. This relief is shown to reduce the crossflow velocity in the boundary layer due to a change in the direction of the near-wall streamlines. On the basis of very simple models based on an analysis of the relief period-averaged boundary layer flow an increase in the Reynolds number of laminar-turbulent transition obtained using the microrelief is estimated. The reasonable surface relief parameters for experimentally investigating the effectiveness of the proposed technique of controlling laminar-turbulent transition are determined. [ABSTRACT FROM AUTHOR]

Highlights • Propose two control method to delay transition over a swept wing using DBD plasma actuators. • The first control method employed one actuators per spanwise fundamental wavelength. These actuators weakened the crossflow directly. • The second control method employed two actuators per spanwise fundamental wavelength. These actuators reduced the crossflow velocity and decreased the growth rates of instability modes. This control method is robust and practical. Abstract This paper investigates the potential use of dielectric-barrier-discharge (DBD) plasma actuators to control the crossflow instability of the boundary layer on a 45° swept NLF-0415 airfoil at − 4 ∘ angle of attack.The disturbance evolution is resolved using nonlinear parabolized stability equations (NPSEs). Actuators are mounted near the leading edge and generate an electric wind perpendicular to the crossflow vortex axes. The spanwise distance between neighboring actuators is first set to be the fundamental wavelength of the unstable mode. This harmonic actuation reduces the disturbance energy only when the induced body force and the disturbance velocity have reverse sign. At certain different locations, the actuators can actually increase the disturbance energy. The robustness in laminar flow control is improved when two actuators per fundamental wavelength are used. This subharmonic actuation leads to reduction in the energy of the primary unstable mode independent of the actuator locations. Finally, the simulation without baseflow forcing reveals that harmonic modes play important roles in this subharmonic control. [ABSTRACT FROM AUTHOR]

The nth-order expansion of the parabolized stability equation (EPSE n) is obtained from the Taylor expansion of the linear parabolized stability equation (LPSE) in the streamwise direction. The EPSE together with the homogeneous boundary conditions forms a local eigenvalue problem, in which the streamwise variations of the mean flow and the disturbance shape function are considered. The first-order EPSE (EPSE1) and the second-order EPSE (EPSE2) are used to study the crossflow instability in the swept NLF(2)-0415 wing boundary layer. The non-parallelism degree of the boundary layer is strong. Compared with the growth rates predicted by the linear stability theory (LST), the results given by the EPSE1 and EPSE2 agree well with those given by the LPSE. In particular, the results given by the EPSE2 are almost the same as those given by the LPSE. The prediction of the EPSE1 is more accurate than the prediction of the LST, and is more efficient than the predictions of the EPSE2 and LPSE. Therefore, the EPSE1 is an efficient e prediction tool for the crossflow instability in swept-wing boundary-layer flows. [ABSTRACT FROM AUTHOR]

In this study, a new centrifugal instability mode, which dominates within the boundary-layer flow over a slender rotating cone, defined by half-angle ψ < 40 ° , is used for the first time to model the problem when an enforced oncoming axial flow is introduced. The resulting similarity solution represents the basic flow more accurately than previous studies in the literature. This mean flow field is subsequently perturbed leading to disturbance equations that are solved via numerical and analytical approaches, importantly yielding favourable comparison with existing experiments. Meanwhile, a formulation consistent with the classic rotating-disk problem has been successful in predicting the stability characteristics of broad rotating cones, defined by half-angle ψ > 40 ° , in axial flow. [ABSTRACT FROM AUTHOR]

The heat addition effect on the stability characteristics of the crossflow in the swept-wing boundary layer is investigated. The physical mechanism of the effect is established. The calculations performed illustrate the variation in the instability mode growth rates under the heating of the surface in a flow and at bulk energy addition. [ABSTRACT FROM AUTHOR]

The crossflow instability of a three-dimensional boundary layer is a main factor affecting the transition around the swept-wing. The three-dimensional boundary layer flow affected by the saturated crossflow vortex is very sensitive to the high frequency disturbances, which foreshadows that the swept wing flow transition will happen. The primary instability of the compressible flow over a swept wing is investigated with nonlinear parabolized stability equations (NPSE). The Floquet theory is then applied to the analysis of the influence of localized steady suction on the secondary instability of crossflow vortex. The results show that suction can significantly suppress the growth of the crossflow mode as well as the secondary instability modes. [ABSTRACT FROM AUTHOR]

The crossflow instability of a three-dimensional (3-D) boundary layer is an important factor which affects the transition over a swept-wing. In this report, the primary instability of the incompressible flow over a swept wing is investigated by solving nonlinear parabolized stability equations (NPSE). The Floquet theory is applied to study the dependence of the secondary and high-frequency instabilities on curvature, Reynolds number and angle of swept (AOS). The computational results show that the curvature in the present case has no significant effect on the secondary instabilities. It is generally believed that the secondary instability growth rate increases with the magnitude of the nonlinear mode of crossflow vortex. But, at a certain state, when the Reynolds number is 3.2 million, we find that the secondary instability growth rate becomes smaller even when the magnitude of the nonlinear mode of the crossflow vortex is larger. The effect of the angle of swept at 35, 45 and 55 degrees, respectively, is also studied in the framework of the secondary linear stability theory. Larger angles of swept tend to decrease the spanwise spacing of the crossflow vortices, which correspondingly helps the stimulation of 'z' mode. [ABSTRACT FROM AUTHOR]