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A series of novel artemisinin (ART) derivatives containing an isothiocyanate (ITC) group were synthesized. All the compounds showed more potent anti-tumor effects than those of parent dihydroartemisinin (DHA) towards glioblastoma multiforme U87 in vitro. Among them, 5b had the strongest cytotoxic activity which exerted its effects in a concentration-dependent but not time-dependent manner (IC50 7.41 μM for 24 h, 7.35 μM for 72 h). Pyknosis was observed in 5b-treated U87 cells. Multiple intrinsic apoptotic pathways were induced by 5b including the upregulation of caspase 9, the release of cytochrome c, an increase of the proapoptotic protein Bax, a decrease of the anti-apoptotic protein Bcl 2, and the activation of execution pathways by the upregulation of caspase 3. In addition to apoptosis, an autophagic mechanism was also involved in 5b-induced cytotoxicity to human GBM U87 cells by upregulating the expression of LC3-II and downregulating p62. Furthermore, 5b also significantly attenuated the migration of U87 cells. Therefore, our results suggest that 5b may be a promising molecule for the further development of a novel drug for the treatment of glioblastoma.

T HE performance of hypersonic transportation vehicles and reentry vehicles is significantly affected by the laminarturbulent transition of boundary-layer flows over vehicle surfaces, as transition has a first-order impact on the lift, drag, stability, control, and surface heating. For a reentry vehicle, transition can lead to an increase in the surface heating rate by a factor of five or more. Hence, the understanding of transition mechanisms and the prediction of transition locations are critical to the development of future hypersonic vehicles [1]. One important area of transition study is the effect of roughness on hypersonic boundary-layer transition. Despite several decades of experimental, theoretical, and numerical studies, the effect of surface roughness on transition is still not fully understood [2].Most previous research has focused mainly on tripping the flow using roughness elements. However, there have been a few reported experimental and numerical studies that demonstrate a delay of transition by roughness elements under certain circumstances. James [3] experimentally studied the effects of Mach number and surface roughness on boundary-layer transition using fin-stabilized hollow tube models in free flights. James found that, for someMach numbers, there exists an optimum roughness height that results in a longer laminar run on a rough surface than a smooth surface. In other words, the roughness element delays the transition process rather than promoting it. Holloway and Sterrett [4] performed a transition experiment in the Langley 20 in. Mach 6 tunnel using a flat plate embedded with spherical roughness elements. They found that, for the cases with the smallest roughness diameters, transition was delayed under certain flow conditions. Although they did not investigate the reasons or conditions behind the delay, they hypothesized that the roughness creates a separated laminar mixing layer that is more stable at higher Mach numbers.More recently, Fujii [5] performed experiments using a 5 deg half-angle sharp cone at Mach 7. The tests were completed at stagnation pressures of 2, 4, and 6 MPa. It was found that, for the higher-pressure cases, the onset of transition was delayed when the wavelength of the wavy wall roughness was roughly equal to that of the unstable second-mode wavelength. Although the delay effect is weak, it is still discernible, and the repeatability of the results is remarkably good. It was speculated that there is a relationship between the wavy wall wavelength and second-mode disturbance that leads to transition. However, the mechanism of the transition delay was unknown and not explored. In addition to experimental testing, several numerical studies have reported the roughness effects on transition delay. Marxen et al. [6] studied the disturbance growth on a flat-plate boundary layer atMach 4.8 with localized two-dimensional (2-D) roughness elements. They found the disturbance was strongly damped downstream of the roughness element around the separation region, which agrees with Holloway and Sterrett’s [4] hypothesis. However, the mechanisms were not investigated. At the same time, Duan et al. [7] from University of California, Los Angeles (UCLA) reported that a 2-D roughness element can damp disturbances if the element is placed downstream of the location where the slow hypersonic boundarylayer mode (mode S) and fast hypersonic boundary-layer mode (mode F) have the same phase velocity (the synchronization location). The details of the work by the UCLA group will be discussed in the next paragraph. Riley et al. [8,9] also numerically studied the stability characteristics of a Mach 4 hypersonic boundary layer over a wedge. On the surface of the wedge, they imposed convex or concave panel buckling (compliant panel) at different locations. They found that, when the panel is placed near the trailing edge of the wedge, the panel can move the boundary-layer transition further downstream. On the other hand, Egorov et al. [10] performed numerical simulations of a Mach 6 supersonic boundary layer over a grooved wavy plate. Their study was motivated by the numerical studies of Balakumar [11] and Egorov et al. [12], which showed the second mode remains neutral in the separated region on a 5.5 deg compression corner. Based on the result of the separated region, the intention of the study by Egorov et al. was to generate short local boundary-layer separations by thewavy wall to decrease disturbance growth. The wavy wall was in the form of nine round arc cavities. It was found that thewavywall design damps a range of high-frequency unstable disturbances that are relevant to the second-mode instability. Bountin et al. [13] later confirmed the results byEgorovet al. [10] that thewavywall damps the unstable secondmode in the high-frequency bandwhile it enhances them at lower frequencies. Their experimental data [13] also showed that the wavy wall damps disturbances, not only at the wavy wall wavelength but also at a wide range of disturbances in different frequencies with different wavelengths. Based on these results, they argued that the stabilization effect of the secondmodeby thewavywall is due to altering themean flow instead of an interference process between the second mode and the wavy wall itself. Since 2009, for the purpose of simulating hypersonic flow with finite height roughness elements, Duan et al. has developed a highorder cut-cell method [14]. The new method was then applied to simulating finite roughness elements in a hypersonic boundary layer at Mach 5.92 [7,14]. Different from the wavy wall idea, as in [5] and [10], they found that the relative location of the 2-D roughness Received 4 December 2014; revision received 7 April 2015; accepted for publication 8 April 2015; published online 8 June 2015. Copyright © 2015 by Kahei Danny Fong. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 1533-385X/15 and $10.00 in correspondence with the CCC. *Graduate Research Assistant, Mechanical and Aerospace Engineering Department. Student Member AIAA. ResearchAssociate,Mechanical and Aerospace EngineeringDepartment. Senior Member AIAA. Professor,Mechanical andAerospace EngineeringDepartment. Associate Fellow AIAA. Graduate Research Assistant, School of Aeronautics and Astronautics. Student Member AIAA. Graduate Student, School of Aeronautics and Astronautics. **Professor, School of Aeronautics and Astronautics. Associate Fellow AIAA.

A scheme is presented to generate atomic entanglement by detecting the transmission spectrum of a coupled-cavity system. In the scheme, two 3-level atoms are trapped in separate cavities coupled by a short optical fiber, and the atomic entanglement could be realized in a heralded way by detecting the transmission spectrum of the coupled-cavity system.

In the authors’ previous paper (Bai and Zhong, J. Comput. Phys. 2019), a very high-order upwind multi-layer compact (MLC) scheme designed to solve smooth multi-scale flow problems with complex physics is presented. In the MLC scheme, the auxiliary equations for the first derivatives are introduced. Accordingly, the first derivatives are evolved simultaneously with the function values and used as additional degrees of freedom in the finite difference approximation. The upwind MLC scheme is subsequently derived on a centered stencil, with an adjustable parameter to introduce small dissipation for stability. It was shown in [1] that the scheme achieved spectral-like resolution within a compact stencil, and it produced very high-order accuracy in flow simulations. However, the MLC scheme shows weak numerical instabilities for a small range of wavenumbers when it is applied to multi-dimensional flows [2], which are mainly triggered by the inconsistency between its one- and two-dimensional formulations. The instability may lead to divergence in long-time multi-dimensional simulations. Moreover, the cross-derivative approximation in the MLC scheme requires an ad-hoc selection of supporting grid points. Also, the cross-derivative approximation is relatively expensive for very high-order cases. In this paper, a new directional multi-layer compact (DMLC) scheme is presented in order to address the remaining problems of the MLC scheme and obtain better performance for multi-dimensional flow simulations. The main idea of the new DMLC scheme is to introduce the auxiliary equation for the cross derivative in multi-dimensional cases, consequently, the spatial discretization can be fulfilled along each dimension independently. With this directional discretization technique, the one-dimensional formulation can be applied to all spatial derivatives in a multi-dimensional governing equation. Therefore, the DMLC scheme overcomes inconsistency between one- and two-dimensional formulations of the original MLC scheme; and it also avoids the ad-hoc cross-derivative approximations. The two-dimensional Fourier analysis demonstrates that all modes of the DMLC scheme are stable for wavenumbers in [0, 2π], and it has better spectral resolution and smaller anisotropic error than the MLC scheme. The stability analysis through matrix method indicates that stable boundary closure schemes are much easier to be obtained in the DMLC scheme. The numerical tests in the linear advection equation and the nonlinear Euler equations validate that the DMLC scheme are more accurate and require less CPU time than the MLC scheme on the same mesh. In particular, the long-time simulation results reveal that the DMLC scheme is always stable for both periodic and non-periodic boundary conditions; while the MLC scheme may diverge in some cases. Overall, the new DMLC scheme shows comprehensive improvements from the original MLC scheme with uniform stability, higher accuracy and spectral resolution, and better computational efficiency for multi-dimensional flows. The improvement is even more significant in very high-order cases which makes the DMLC scheme appropriate for solving complex flow problems.

This paper presents a numerical-simulation study of transient flow over a blunt compression cone under the effect of freestream hotspot perturbations. This study is motivated by concurrent wind-tunnel laser-spot experiments carried out at Purdue University. The flow conditions used in the simulation are based on the experimental conditions. The simulation is performed using a high-order shock-fitting finite-difference scheme. The simulation results show that the hotspot is able to excite second-mode instability, where the instability growth is found to be dominant in the boundary layer. The receptivity mechanism is investigated by comparing the simulated results with linear-stability theory. Fast acoustic waves generated by hotspot–shock interaction excite the boundary-layer disturbances. Also, the synchronization of mode F and mode S leads to the dominance of boundary-layer disturbances by the growing second mode.

A new high-order shock-fitting method with thermochemical nonequilibrium and finite-rate chemistry boundary conditions for graphite ablation is presented. The method is suitable for direct numerical simulation of boundary-layer transition in a hypersonic real-gas flow with graphite ablation. The new method is validated by comparison with three computational datasets and one set of experimental data. Direct numerical simulations were run for a 7 deg half-angle blunt cone at Mach 15.99 to find how graphite ablation and thermochemical nonequilibrium affect boundary-layer receptivity and instability. The real-gas simulation is compared with ideal-gas simulations that set their wall temperature and wall blowing from the real-gas simulation. Weak planar fast-acoustic waves in the freestream are used to perturb the steady base flow. A 525 kHz second-mode wave was found to be significantly unstable for the real-gas simulation, whereas in the ideal-gas simulations, no significant flow instability was seen. For the...

A new high-order cut-cell method is used to numerically simulate two dimensional roughness effects on modal growth in a hypersonic boundary layer. The steady base flow is simulated first by solving the compressible Navier–Stokes equations. Perturbations correspond to Mode S wave at 100 kHz and a wall normal velocity pulse with a continuous frequency spectrum up to 1 MHz are imposed onto the mean flow with roughness separately. A Fast Fourier Transform (FFT) is used to decompose the perturbation and to study the evolution of perturbations at different frequencies. The number of roughness elements in the boundary layer is changed for two test cases. Moreover, the effect of roughness height is studied in the Mode S simulation. It is found that when the roughness element is located at the synchronization point of a particular frequency (i.e. synchronization frequency), perturbations at frequencies higher than the synchronization frequency are damped. On the other hand, perturbations at frequencies lower than the synchronization frequency are amplified by the roughness. The same conclusion is drawn in cases where the roughness location is changed. The relation between roughness and the synchronization point is a candidate to explain roughness delayed transition seen experimentally.

There has been renewed interest in supersonic modes in hypersonic boundary layers, which have been previously thought to be insignificant due to their smaller amplitudes than Mack’s traditional second mode. Supersonic modes are associated with an unstable second mode synchronizing with the slow acoustic spectrum, causing sound to radiate outwards from the boundary layer. Because supersonic modes have not been observed experimentally, the majority of previous investigations either relied on Linear Stability Theory (LST) to study supersonic modes on a flat plate or observed them in the context of other research objectives. This two-part study uses a combined LST and Direct Numerical Simulation (DNS) approach to investigate the mechanism of supersonic modes in Mach 5 flow over a blunt cold-wall cone with thermochemical nonequilibrium effects. Paper I focuses on LST with new shock boundary conditions, whereas Paper II [C.P. Knisely and X. Zhong, “Sound radiation by supersonic unstable modes in hypersonic blunt cone boundary layers. II. Direct numerical simulation,” Phys. Fluids 31, 024104 (2019)] focuses on DNS with the overall goal of investigating the impact of supersonic modes on transition. LST results indicate that supersonic modes exist in the flow with wall-to-free-stream temperature ratio Tw/T∞ = 0.2 and create an abnormal growth pattern. However, supersonic modes were not shown to exist using LST in the case with Tw/T∞ = 0.667. Subsequent DNS analysis in Paper II shows supersonic modes in the Tw/T∞ = 0.667 case, although they are significantly weaker than the second mode and are unlikely to lead to transition. Understanding the mechanism of supersonic modes can yield more accurate transition location predictions leading to improved estimates for drag and heat transfer to the vehicle.

Supersonic modes, previously thought to be insignificant due to their smaller amplitude than Mack’s traditional second mode, occur in hypersonic boundary layers when a disturbance travels supersonically with respect to the mean flow outside the boundary layer, causing outward-radiating acoustic waves. Very few previous studies perform Direct Numerical Simulation (DNS) of supersonic modes and instead rely on Linear Stability Theory (LST). This combined LST and DNS study investigates supersonic modes in Mach 5 flow over a blunt cold-wall cone. An LST analysis was performed in Paper I [C. P. Knisely and X. Zhong, “Sound radiation by supersonic unstable modes in hypersonic blunt cone boundary layers. I. Linear stability theory,” Phys. Fluids 31, 024103 (2019)], whereas DNS is the focus of Paper II. The overall goal is to determine the mechanism of supersonic modes and the conditions under which they exist. Compared to previous pure LST studies, DNS provides the advantage of making fewer limiting assumptions and can resolve interactions between modes. The results here indicate the excitation of supersonic modes via modal interactions not resolved with LST, suggesting the inadequacy of pure LST analyses concerning supersonic modes. Unsteady DNS results verified supersonic modes in the flow with wall-to-free-stream temperature ratio Tw/T∞ = 0.2, lending credence to the modes’ physical existence. However in the case of Tw/T∞ = 0.667, sound radiation was also found in DNS while LST predicted a stable supersonic mode. The mechanism for supersonic modes is attributed to a modal interaction between mode F1, mode S, and the slow acoustic spectrum. Therefore, it is necessary to perform combined LST and DNS studies of supersonic modes to reliably predict their presence and impact on transition to turbulence.

Y 2 O 3 :Eu 3+ nanophosphors were fabricated by direct thermal decomposition of monohydrate yttrium europium propionates, which can be facilely obtained via cocrystallization. The composition and structure of the cocrystallized product were confirmed by X-ray diffractionmetry (XRD), simultaneous differential scanning calorimetry (DSC) and thermogravimetric analysis (TGA) as well as inductively coupled plasma mass (ICP-MS) spectrometry. The effects of decomposition temperature on the resultant microstructures such as crystallinity, particle size and particle dispersibility were investigated by XRD and transmission electron microscopy (TEM). The dependence of their luminescent properties on decomposition temperature was also studied by fluorescence spectrophotometry. The results showed that the microstructures and luminescent intensities of Y 2 O 3 :Eu 3+ nanophosphors could be tuned by varying calcination temperatures of cocrystallized yttrium europium propionates from 600 to 1000 °C.

THE performance of hypersonic transportation vehicles and reentry vehicles is significantly affected by the laminar-turbulent transition of boundary-layer flows over vehicle surfaces as transition has a first-order impact on lift, drag, stability, control, and surface heating of these vehicles. For example, the roughness-induced transition is an important consideration in the design of thermal protection systems [1,2]. For a reentry vehicle transition can lead to an increase in surface-heating rate by a factor of five or more. Hence, the understanding of transition mechanisms is critical to the development of future hypersonic vehicles [3]. The dominant mechanism in the roughness-induced transition is the transient growth associated with purely wall-normal and spanwise velocity perturbations, which finally evolve into streak-like motion. Transient growth has been studied by many researchers. Andersson et al. [4] calculated the transient growth of a flat-plate boundary layer to steady disturbances and found that the maximum transient growth scales linearly with the distance from the leading edge, which was consistent with the results of Hanifi et al. [5]. Collis and Lele [6] investigated the stationary crossflow vortices in a threedimensional boundary layer over a swept wing due to surface roughness near the leading edge. The results showed that the initial amplitude of crossflow vortices is enhanced by convex surface curvature and strongly reduced by nonparallel flow effects. White and Ergin [7] studied the transient growth of a Blasius boundary layer generated by a spanwise array of roughness elements. The results indicated that the energy of the roughness-induced disturbances is proportional to the roughness-height-based Reynolds number.White et al. [8] further investigated the effects of the height and diameter of cylindrical roughness element on transient growth. Their results showed that the energy of transient growth is proportional to the square of the roughness-height-based Reynolds number. In addition, transient growth strongly depends on roughness diameter. Choudhari and Fischer [9,10] examined the transient growth in a laminar boundary layer due to a spanwise array of circular disks at the surface. The effects of roughness height, size, and shapewere explored. Their results indicated that the energy of transient growth is consistent with the scale of White et al. [8]. One explanation to the roughness-induced transition is the transient growth theory pioneered by Reshotko and Tumin [11] where the roughness-induced velocity perturbations are assumed to be proportional to roughness height. Wang and Zhong [12] studied the transient growth of a flat-plate boundary layer to small surface roughness periodic in the spanwise direction. The results showed that the assumption, the roughness-induced velocity perturbations are proportional to roughness height, is only valid when roughness height is smaller than approximately one thirty-fifth of the local boundary-layer thickness. Therefore, the transient growth theory cannot be directly applied to the transient growth induced by finite-height roughness [8–10]. Unfortunately, the roughness-induced transition in hypersonic boundary layers, especially that induced by arbitrary surface roughness, is still poorly understood due to the limitations in experimental facilities and numerical methods [13]. Marxen et al. [14] studied the effect of a two-dimensional (2-D) roughness element on the stability of a hypersonic boundary layer. They showed that in the vicinity of the separation regions located near the edges of the roughness the destabilization of a second-mode disturbance occurs for a certain frequency. Even though roughness is generally expected to lead to an early transition several past studies have shown that roughness can stabilize boundary-layer flows and delay transition in some situations. For example, Holloway and Sterrett [15] concluded from their experiments that surface roughness can delay transition when its height is less than the local boundary-layer thickness. Fujii [16] found that a wavy roughness can suppress the transition of a Mach 7.1 boundary layer. Although no physical explanation is given the repeatable results indicate that some unknown mechanisms must have been presented. Recently, Wang et al. [17] studied the response of a Mach 8 flow over a 5.3 deg half-angle wedge to wall blowing– suction and showed that the synchronization point plays a critical role in the stability of hypersonic boundary layers. The results of Wang et al. [17] indicate that the location of roughness with respect to the synchronization point might play a role in the roughness-caused second-mode destabilization [14] and transition delay [15,16]. In this Technical Note, we present our numerical simulation results on the stabilization of aMach 5.92 flow over a flat plate by 2-D finiteheight roughness. The main objective is to investigate the effect of roughness location on mode S instability. Although three-dimensional surface roughness is more relevant to the roughness-induced transition the current numerical simulations on 2-D finite-height roughness are still valuable for the effect of roughness on the stability of hypersonic boundary layers. Figure 1 shows a schematic of the problem where boundary-layer waves are excited by a blowingsuction slot near the leading edge and a finite-height roughness is put downstream of the slot. Specifically, four cases of numerical simulations are conducted with the roughness at different locations. The results show that mode S is destabilized only when the roughness element is located closely upstream of the synchronization point. When the roughness element is downstream of the synchronization Presented at the 47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition, Orlando, Florida, 5–8 January 2009; received 17 October 2011; revision received 13 June 2012; accepted for publication 14 June 2012; published online 21 November 2012. Copyright © 2012 by XiaowenWang. Published by the American Institute of Aeronautics andAstronautics, Inc., with permission. Copies of this papermay be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 1533-385X/12 and $10.00 in correspondence with the CCC. *Graduate Research Assistant, Mechanical and Aerospace Engineering Department. Student Member AIAA. ResearchAssociate,Mechanical and Aerospace EngineeringDepartment. Senior Member AIAA. Professor,Mechanical andAerospace EngineeringDepartment. Associate Fellow AIAA.

This paper presents experimental and theoretical investigations into excited-state intramolecular proton transfer (ESIPT) in new chromophores with hydroxyl and imino groups under one- and two-photon excitation. The results show that internal hydrogen bonding exhibits a remarkable influence on the maximum absorption wavelength of 2-[(4'-N,N-diethylaminodiphenylethylene-4-ylimino)methyl]phenol (C1) and 2-[(4'-methoxyl-diphenylethylene-4-ylimino)methyl]phenol (C3). Compounds C1 and C3 exhibit well-separated dual fluorescence emission bands under one- and two-photon excitation. The second fluorescence peaks of C1 and C3 are characterized by much larger Stokes shift than the first normal peaks (ca. 140 vs. 30 nm). 4-[(4'-N,N-Diethylaminodiphenylethylene-4-ylimino)methyl]phenol (C2) and 4-[(4'-methoxyldiphenylethylene-4-ylimino)methyl]phenol (C4) display single emission bands with small Stokes shifts (ca. 30 nm) in various solvents under one- and two-photon excitation. Furthermore, the first emission maxima of C1 and C3 are almost identical to the maximum fluorescence emission wavelengths of C2 and C4, respectively. These results show that C1 and C3 can undergo ESIPT via a reasonable six-membered ring, while there is no ESIPT in C2 and C4 under one- and two-photon excitation. Compounds C1 and C2 have larger two-photon absorption cross-sections under various near-infrared laser frequencies tuned from 700 to 880 nm. Molecular geometry optimization of the phototautomers (enol and keto) was performed to analyze the experimental results. The possibility of using these chromophores for metal ions as chemosensors of was thoroughly investigated. In DMF C3 exhibits excellent sensing responses to Zn(2+) and Fe(3+) ions through a greatly increased greatly and a largely reduced emission, respectively. In methanol disappearance of ESIPT emission with added Zn(2+) ions confirms its existence. The binding constants of C3 with Zn(2+) and Fe(3+) ions in DMF are also estimated.

The prediction of the laminar-turbulent transition of boundary layers is critically important to the development of hypersonic vehicles because the transition has a first-order impact on aerodynamic heating, drag, and vehicle operation. The success of transition prediction relies on a fundamental understanding of the relevant physical mechanisms. In the 20 years since the review by Kleiser & Zang (1991) on the direct numerical simulation (DNS) of the boundary-layer transition, significant progress has been made on DNS in the hypersonic flow regime and in the spatial DNS approach. Many high-order shock-capturing and shock-fitting finite-difference methods have been developed and extensively applied to numerical simulations of the hypersonic boundary-layer transition. DNS has become a powerful research tool and has led to discoveries of new transition mechanisms. This article reviews the recent progress of DNS on hypersonic boundary-layer receptivity, instability, and transition. The current status and future directions are also presented.

conduct a linear stability analysis on Stetson andRushton’s Mach 5.5experiments in which the reversalis observed. Three cones with nose radiiof 0.156, 0.5,and 1.5 in.,covering both the small andlarge bluntness regions, are used to study the effect of nose bluntness on stability and transition. It is found that, if only the second-mode instabilities are considered, the onset of instability is always delayed as the nose bluntness increases. The linear stability theory calculations show no reversal on the growth of the second-mode instability.

In the laminar-turbulent transition of boundary-layer flows, the spatial development of boundary-layer waves, especially that of unstable waves, is important and indispensable. In this paper, the response of a Mach 8 flow over a 5.3° half-angle sharp wedge to wall blowing―suction is studied by numerical simulations. Steady base flow is obtained by solving the compressible Navier―Stokes equations with a combination of a fifth-order shock-fitting method and a second-order total-variation-diminishing scheme. In stability simulations, wall blowing―suction is introduced through an actuator on the wedge surface. The unsteady flow simulation is carried out using the shock-fitting method. The results show that mode F, mode S, acoustic waves, and entropy/vorticity waves are simultaneously excited by wall blowing―suction and coexist in the boundary layer just downstream of the actuator. For blowing―suction at a specific frequency, the results also show that mode S is strongly excited when the actuator is located upstream of the corresponding synchronization point. There is no significant amplification of pressure perturbation when the actuator is downstream of the synchronization point. This result represents a mixture of the receptivity and the downstream growth of mode S. The exact cause and mechanism of this result are not clear. However, such a result is obtained for wall blowing―suction at all frequencies considered in the current study. To excite a strong mode S at a specific frequency, the result indicates that it is necessary to place the blowing―suction actuator upstream of the corresponding synchronization point. Further theoretical analysis is needed to reveal the mechanism behind the numerical results.

Flower-like CuS nanostructures have been synthesized via a liquid precipitation route by the reaction between CuCl2·2H2O and thioacetamide (CH3CSNH2, TAA) in the ionic liquid 1-butyl-3-methyl imidazole six hexafluorophosphoric acid salts ([BMIM][PF6]) aqueous solution at room temperature. The products were characterized by X-ray powder diffraction (XRD), field emission scanning electronic microscopy (FESEM), Brunauer-Emmett-Teller (BET), Ultraviolet-Visible Spectrophotometer (UV-Vis) and Photoluminescence (PL) techniques. The as-prepared CuS nanostructures have a mean diameter of about 1 μm. A plausible mechanism was proposed to explain the formation of CuS nanostructures. The effects of experimental parameters on the formation of the products were also explored. With BET theory, it is found that the as-prepared CuS nanostructures have a specific area of 39m2/g. The Barrett-Joyner-Halenda (BJH) pore size distribution of the as-prepared CuS nanostructures presents smaller pores centers about 60 nm.The UV-Vis and PL curves indicate that the asprepared CuS nanostructures are promising candidates for the development of photoelectric devices.

Direct numerical simulations of receptivity in a boundary layer over a flat plate and a sharp wedge were carried out with two-dimensional perturbations introduced into the flow by periodic-in-time blowing-suction through a slot. The free stream Mach numbers are equal to 5.92 and 8 in the cases of adiabatic flat plate and sharp wedge, respectively. The perturbation flow field was decomposed into normal modes with the help of the multimode decomposition technique based on the spatial biorthogonal eigenfunction system. The decomposition allows filtering out the stable and unstable modes hidden behind perturbations having another physical nature.

Laminar-turbulent transition of hypersonic boundary layers can be affected significantly by the existence of surface roughness. Currently many important mechanisms of roughness-induced transition are not well understood. In recent years, direct numerical simulation (DNS) has been extensively applied for investigating instability and transition mechanisms of hypersonic boundary layers. Most of the past DNS studies, however, have been based on body-fitted grids for smooth surfaces without roughness. Due to complex geometry of embedded roughness, the use of body-fitted grids can be very difficult for flow with arbitrary surface roughness. In this paper, we present a new high-order cut-cell method to overcome the natural complexities in grid generation around arbitrary surface of roughness. The new method combines a non-uniform-grid finite-difference method for discrete grid points near the solid boundary and a shock-fitting method for the treatment of the bow shock. The non-uniform-grid finite-difference formulas are expressed in a general explicit form so that they can be applied to different multi-dimensional problems without any modification. The computational accuracy of new algorithms of up to O(h^4) are tested on several one- and two-dimensional elliptic equations in irregular domains. In addition, the new method is applied to the simulation of the receptivity process of a Mach 5.92 flow over a flat plate under the combined effect of an isolated surface roughness element and surface blow and suction. A good agreement is found between the unsteady flow results and those obtained by a Linear Stability Theory (LST).

High-order methods that can resolve interactions of flow-disturbances with shock waves are critical for reliable numerical simulation of shock wave and turbulence interaction. Such problems are not well understood due to the limitations of numerical methods. Most of the popular shock-capturing methods are only first-order accurate at the shock and may incur spurious numerical oscillations near the shock. Shock-fitting algorithms have been proposed as an alternative which can achieve uniform high-order accuracy and can avoid possible spurious oscillations incurred in shock-capturing methods by treating shocks as sharp interfaces. We explore two ways for shock-fitting: conventional moving grid set-up and a new fixed grid set-up with front tracking. In the conventional shock-fitting method, a moving grid is fitted to the shock whereas in the newly developed fixed grid set-up the shock front is tracked using Lagrangian points and is free to move across the underlying fixed grid. Different implementations of shock-fitting methods have been published in the literature. However, uniform high-order accuracy of various shock-fitting methods has not been systematically established. In this paper, we carry out a rigorous grid-convergence analysis on different variations of shock-fitting methods with both moving and fixed grids. These shock-fitting methods consist of different combinations of numerical methods for computing flow away from the shock and those for computing the shock movement. Specifically, we consider fifth-order upwind finite-difference scheme and shock-capturing WENO schemes with conventional shock-fitting and show that a fifth-order convergence is indeed achieved for a canonical one-dimensional shock-entropy wave interaction problem. We also show that the method of finding shock velocity from one characteristic relation and Rankine-Hugoniot jump condition performs better than the other methods of computing shock velocities. A high-order front-tracking implementation of shock-fitting is also presented in this paper and nominal rate of convergence is shown. The front-tracking results are validated by comparing to results from the conventional shock-fitting method and a linear-interaction analysis for a two-dimensional shock disturbance interaction problem.

In this paper, we described the synthesis and characterization of new diphenylethylene bearing imino group. We concentrated particularly on the investigation of the possibility of the excited state intramolecular charge transfer (ESIPT) of the new dyes experimentally and theoretically. The absorption and fluorescence spectroscopy of the dyes were determined in various solvents. The results showed that the maximal absorption wavelength of 2-[(4′-N,N-dimethylamino-diphenylethylene-4-ylimino)methyl]phenol (C1) and 4-[(4′-N,N-dimethylamino-diphenylethylene-4-ylimino)methyl]phenol (C2) exhibited almost independence on the solvent polarity. While as contrast, the maximal fluorescence wavelength of the dyes showed somewhat dependence on the solvent polarity. In particular, C1 displayed well-separated dual fluorescence spectroscopy. The second fluorescence peak was characterized with an "abnormal" fluorescence emission wavelength in aprotic solvents with large Stokes shift (ca. 140 nm in THF), which was much more than normal Stokes shift (ca. 30 nm in THF). This emission spectroscopy could be assigned to ESIPT emission. On the other hand, the ESIPT fluorescence of C1 was much reduced or lost in the protic solvents. While, only normal fluorescence emission was detected in various solvents. Although the absorption maxima of C1 exhibited about 10 nm red-shift with respect to those of C2, the normal fluorescence maxima of C1 and C2 were almost identical in various solvents. These results suggested that C1 could undergo ESIPT, but C2 was not able to proceed ESIPT. The molecular geometry optimization of phototautomers in the ground electronic state (S0) was carried out with HF method (Hartree-Fock) and at DFT level (Density Functional Theory) using B3LYP both, while the CIS was employed to optimize the geometries of the first singlet excited state (S1) of the phototautomers of C1 and C2 respectively. The properties of the ground state and the excited state of the phototautomers of C1 and C2, including the geometrical parameter, the energy, the frontier orbits, the Mulliken charge and the dipole moment change were performed and compared completely. The data were analyzed further based on our experimental results. Furthermore, the absorption and fluorescence spectra were calculated in theory and compared with the measured ones. The rate constant of internal proton transfer (9.831×1011 s−1) of C1 was much lower than that of salicylidene methylamine (C3, 2.045×1015 s−1), which was a typical Schiff base compound and was well demonstrated to undergo ESIPT easily under photoexcitation.

Flows in which shock waves and turbulence are present and interact dynamically occur in a wide range of applications, including inertial confinement fusion, supernovae explosion, and scramjet propulsion. Accurate simulations of such problems are challenging because of the contradictory requirements of numerical methods used to simulate turbulence, which must minimize any numerical dissipation that would otherwise overwhelm the small scales, and shock-capturing schemes, which introduce numerical dissipation to stabilize the solution. The objective of the present work is to evaluate the performance of several numerical methods capable of simultaneously handling turbulence and shock waves. A comprehensive range of high-resolution methods (WENO, hybrid WENO/central difference, artificial diffusivity, adaptive characteristic-based filter, and shock fitting) and suite of test cases (Taylor-Green vortex, Shu-Osher problem, shock-vorticity/entropy wave interaction, Noh problem, compressible isotropic turbulence) relevant to problems with shocks and turbulence are considered. The results indicate that the WENO methods provide sharp shock profiles, but overwhelm the physical dissipation. The hybrid method is minimally dissipative and leads to sharp shocks and well-resolved broadband turbulence, but relies on an appropriate shock sensor. Artificial diffusivity methods in which the artificial bulk viscosity is based on the magnitude of the strain-rate tensor resolve vortical structures well but damp dilatational modes in compressible turbulence; dilatation-based artificial bulk viscosity methods significantly improve this behavior. For well-defined shocks, the shock fitting approach yields good results.

The laminar–turbulent transition of hypersonic boundary layers has a significant effect on drag calculation and aerothermal design of hypersonic vehicles. Recent research has shown that one possible explanation to roughnessinduced bypass transition is transient growth theory. However, it is not known how to generate the optimal disturbances computed by transient growth theory. Furthermore, there has not been any direct numerical simulation study on transient growth in hypersonic boundary layers. The objectives of this paper are to study the receptivity of a Mach 5.92 flow over a flat plate to three-dimensional surface roughness and the effect of spanwise wave number on the receptivity. Steady base flow is computed by solving compressible Navier–Stokes equationswith a combination of a fifth-order shock-fitting method and a second-order total variation diminishing scheme. In receptivity simulations, small surface roughness is introduced on the plate. The numerical results show that counterrotating streamwise vortices and transient growth are induced by surface roughness, however, transient growth is generally weak due to the small height of roughness. For the six cases considered, surface roughness with the spanwise wave number being 0.0101 has the strongest excitation of transient growth.

For simultaneous localization and mapping (SLAM) of mobile robots, an innovative solution is proposed, named adaptive square root cubature Kalman filter based SLAM algorithm (ASRCKF-SLAM). The main contribution of the proposed algorithm lies that: 1) Square root factors are used in the proposed ASRCKF-SLAM algorithm to improve the calculation efficiency by avoiding the time-consuming Cholesky decompositions. 2) Using the adaptive Sage-Husa estimator to solve the large estimation errors or even divergence problem caused by the time-varying or unknown noise. Simulation results obtained demonstrate that the proposed ASRCKF-SLAM algorithm is superior to the existed SLAM method in the aspect of estimation accuracy and computational efficiency.

Experimental and numerical results have shown that two-dimensional surface roughness can stabilize a hypersonic boundary layer dominated by second-mode instability. It is sought to understand how this physical phenomenon extends from an airflow under a perfect gas assumption to that of a real gas. To these ends, a new high-order shock-fitting method that includes thermochemical nonequilibrium and a cut-cell method to handle complex geometries unsuitable for structured body-fitted grids is presented. The new method is designed specifically for direct numerical simulation of hypersonic boundarylayer transition in a hypersonic real-gas flow with arbitrary shaped surface roughness. The new method is validated and shown to perform comparably to a high-order method with a body-fitted grid. For a Mach 10 flow over a flat plate with a real-gas model, a two-dimensional roughness element was found to stabilize the second mode when placed downstream of the synchronization location which is consistent with previous research for perfect-gas flows. For a Mach 15 flow over a flat plate, a two-dimensional surface roughness element stabilizes the second-mode instability more effectively in a real gas than in a perfect gas.

In this paper simple polynomial interpolation is used to derive arbitrarily high-order compact schemes for the first derivative and tridiagonal compact schemes for the second derivative (consisting of three second derivative nodes in the interior and two on the boundary) on non-uniform grids. Boundary and near boundary schemes of the same order as the interior are also developed using polynomial interpolation and for a general compact scheme on a non-uniform grid it is shown that polynomial interpolation is more efficient than the conventional method of undetermined coefficients for finding coefficients of the scheme. The high-order non-uniform schemes along with boundary closure of up to 14th order thus obtained are shown to be stable on a non-uniform grid with appropriate stretching so that more grid points are clustered near the boundary. The stability and resolution properties of the high-order non-uniform grid schemes are studied and the results of three numerical tests on stability and accuracy properties are also presented.

In the numerical simulation of transient reacting flow, standard explicit calculation is prohibitively expensive because of the small time steps needed to address the stiffness of a governing differential system. To circumvent this, new hybrid implicit‐explicit methods proposed treat the stiffness, whereas the underlying time-step control is governed by the Courant stability criterion. Because the coefficients of both explicit and implicit operations are entirely determined by solving the necessary conditions of accuracy and L stability without any extra assumptions, the methods are more generalized than other similar methods in the literature. Two families of semi-implicit Runge‐ Kutta schemes are developed for split differential equations in the form of u � = f (t, u) + g(t, u), where f is treated explicitly and g is simultaneously treated implicitly. Like the rest of all the developed schemes, a low-storage family of semi-implicit schemes is also derived to be globally high-order accurate and L stable for implicit calculations. In a companion paper (Yoh, J. J., and Zhong, X., “New Hybrid Runge‐Kutta Methods for Unsteady Reactive Flow Simulation: Applications,” AIAA Journal ,V ol. 42, No. 8, 2004, pp. 1601‐1611) the new schemes are tested to solve a wide range of applications in high-speed flow physics involving combustion.

The applicability is examined of the new hybrid Runge-Kutta methods derived in the companion paper by specifically analyzing both nonstiff and stiff system of equations that represent multidimensional reactive flows. In a reactive flow simulation, the standard explicit calculation is prohibitively expensive because of the small time steps needed to address the stiffness of a governing differential system. The new hybrid Runge-Kutta schemes are suitable for this task because the time-step size is controlled by the Courant condition, whereas the stiffness is treated by an unconditionally stable method. This set of methods also meets the modern computing needs of high-order accuracy and low-storage requirement. A representative scheme is used to simulate a series of combustion problems that include model equations with stiff sources and multidimensional detonations with complex chemical kinetics

The direct numerical simulation of receptivity, instability and transition of hypersonic boundary layers requires high-order accurate schemes because lower-order schemes do not have an adequate accuracy level to compute the large range of time and length scales in such flow fields. The main limiting factor in the application of high-order schemes to practical boundary-layer flow problems is the numerical instability of high-order boundary closure schemes on the wall. This paper presents a family of high-order non-uniform grid finite difference schemes with stable boundary closures for the direct numerical simulation of hypersonic boundary-layer transition. By using an appropriate grid stretching, and clustering grid points near the boundary, high-order schemes with stable boundary closures can be obtained. The order of the schemes ranges from first-order at the lowest, to the global spectral collocation method at the highest. The accuracy and stability of the new high-order numerical schemes is tested by numerical simulations of the linear wave equation and two-dimensional incompressible flat plate boundary layer flows. The high-order non-uniform-grid schemes (up to the 11th-order) are subsequently applied for the simulation of the receptivity of a hypersonic boundary layer to free stream disturbances over a blunt leading edge. The steady and unsteady results show that the new high-order schemes are stable and are able to produce high accuracy for computations of the nonlinear two-dimensional Navier-Stokes equations for the wall bounded supersonic flow.

In this paper, we continue to study the mechanisms of the receptivity of the supersonic boundary layer to free-stream disturbances by using both direct numerical simulation and linear stability theory. Specifically, the receptivity of a Mach 4.5 flow over a flat plate to free-stream fast acoustic waves is studied. The receptivity to free-stream slow acoustic waves, entropy waves and vorticity waves will be studied in the future. The oblique shock wave induced by the boundary-layer displacement plays an important role in the receptivity because the free-stream disturbance waves first pass through the shock before entering the boundary layer. A high-order shock-fitting scheme is used in the numerical simulations in order to account for the effects of interactions between free-stream disturbance waves and the oblique shock wave. The results show that the receptivity of the flat-plate boundary layer to free-stream fast acoustic waves leads to the excitation of both Mack modes and a family of stable modes, i.e. mode I, mode II, etc. It is found that the forcing fast acoustic waves do not interact directly with the unstable Mack modes. Instead, the stable mode I waves play an important role in the receptivity process because they interact with both the forcing acoustic waves and the unstable Mack-mode waves. Through the interactions, the stable mode I waves transfer wave energy from the forcing fast acoustic waves to the second Mack-mode waves. The effects of incident wave angles, forcing wave frequencies, and wall temperature perturbation conditions on the receptivity are studied. The results show that the receptivity mechanisms of the second mode are very different from those of modes I and II, which leads to very different receptivity properties of these discrete wave modes to free-stream fast acoustic waves with different incident wave angles, frequencies, and different wall boundary conditions. The maximum receptivities of the second mode, mode I and mode II to planar free-stream fast acoustic waves are obtained when incident wave angles approximately equal 26 ◦ ,4 5 ◦ , and 18 ◦ ,r espectively. The results of receptivity to a beam of free-stream fast acoustic waves show that the leading edge is one of the most efficient regions for receptivity.

This paper is the first part of a two-part study on the mechanisms of the receptivity to disturbances of a Mach 4.5 flow over a flat plate by using both direct numerical simulations (DNS) and linear stability theory (LST). The main objective of the current paper is to study the linear stability characteristics of the boundary-layer wave modes and their mutual resonant interactions. The numerical solutions of both steady base flow and unsteady flow induced by forcing disturbances are obtained by using a fifth-order shock-fitting method. Meanwhile, the LST results are used to study the supersonic boundary-layer stability characteristics relevant to the receptivity study. It is found that, in addition to the conventional first and second modes, there exist a family of stable wave modes in the supersonic boundary layer. These modes play a very important role in the receptivity process of excitation of the unstable Mack modes, especially the second mode. These stable modes are termed mode I, mode II, etc., in this paper. Though mode I and mode II waves are linearly stable, they can have resonant (synchronization) interactions with both acoustic waves and the Mack-mode waves. Therefore, the stable wave modes such as mode I and mode II are critical in transferring wave energy between the acoustic waves and the unstable second mode. The effects of frequencies and wall boundary conditions for the temperature perturbations on the boundary-layer stability and receptivity are also studied.

This paper investigates, by numerical simulation, the effects of an imposed magnetic field on a weakly ionized Mach 4.5 boundary layer. The main emphasis of the study is on magnetohydrodynamic (MHD) effects on the second mode instability in supersonic boundary layer. The imposed magnetic fields are generated by placing two-dimensional magnetic dipoles below the flat plate surface. The gas is assumed to have a constant electrical conductivity of 100 mho/m. The magnetic Reynolds number of the flow is small so that the induced magnetic field in the flow is neglected. The governing equations of the MHD flow, which are the Navier–Stokes equations with the applied magnetic force terms, are computed by a fifth-order shock-fitting numerical scheme. A series of cases with different imposed magnetic fields have been investigated on the influences of imposed magnetic field on both the mean flow and on the second mode stability. It is found that the imposed magnetic fields significantly retard the streamwise velocity and reduce the local skin friction in the mean flow. For the case of a strong imposed magnetic field, a local separation region is generated in the mean flow with a strong adverse pressure gradient. Meanwhile, the second mode wave disturbances are found to be stabilized by the imposed magnetic fields, even for the case with strong adverse pressure gradient and a local separated flow region. This strong overall stabilization of the second mode wave is believed to be caused by the alteration of the steady base flow by the magnetic field. The results presented in this paper are the first concrete results on the interaction of second instability mode with magnetic field in a supersonic boundary layer.

Introduction T HE calculation of aerodynamic heating rates on hypersonic vehicles remains a challenging problem. Such analysis often involvesmultidimensionalgeometrieswith leading-edgebluntness and fuselage/wing combinations of arbitrary cross sections. In addition, hypersonic  ight conditions must take into account the dissociation of air at high temperatures, as well as the resulting effects on surfaceheating.Numerical simulation of full compressible Navier–Stokes equations for three-dimensional  owŽ elds around these vehicles is still currently too computationallyexpensive to be routinely used in practical aerodynamic heating calculations. Becauseof the large computationaltime required to compute viscous  ows over full aircraft geometries, inviscid  ow solutions are computed instead. If an inviscid ow solutionis known, then a practical engineering approach for aerodynamic heating calculations is to use a boundary-layerapproximation.Current aerodynamic heatingmethods that are used in thisNote forhypersonic ow conditions were developed at NASA Dryden Flight Research Center and are summarized by Quinn and Gong.1 They provide heating models based on laminar and turbulent  ow, using a constant or variable entropy assumption, with the ability to compute transient heating rates as well as steady heating rates. The motivation for this study is the need for accurate prediction of aerodynamicheating aroundhypersonicvehicles.NASA Dryden Flight Research Center is interested in extending their hypersonic heatingcalculationcodewith advancedcomputational uid dynamics (CFD) inviscid solutions.NASA Dryden Flight ResearchCenter has the requirementto conduct ight researchon high-speedaircraft. To accomplish this mission and ensure  ight safety, it is necessary to have the capability of predicting transient surface heating rates and temperatures. The accurate calculation of surface heating rates on hypersonic vehicles is dependent on an accurate inviscid CFD numerical solution of the mean aerodynamic  ow properties. The shape of the hypersonicvehiclemay vary from a simple cone to a complex threedimensional aircraft. It is because of this constraint that the shock

Recent research has shown that 2-D roughness of finite height can damp the second mode instability under certain conditions. For instance, it has been shown that the relative location of 2-D roughness element and the synchronization point is important in determining the 2-D roughness effect on modal growth. It was found that if a roughness element is placed downstream of a disturbance's synchronization point, the disturbance is damped. In this paper, an extensive DNS parametric study on the effect of second mode instability from roughness height, width and spacing between roughness elements are conducted. The results reveal that a taller roughness within one boundary layer thickness height results a stronger damping effect on the second mode. On the other hand, the roughness width effect is relatively insignificant compared with the height effect. Lastly, the spacing between roughness elements has been studied. The results show the increasing in spacing of roughness elements has both advantages and disadvantages. For the second mode control using 2-D roughness elements, roughness elements spacing of 10 times the roughness width has been shown to have the best overall damping effect in a broad frequency bandwidth. Overall, our extensive study has shown that roughness height is an important factor for the amplification/damping effect. For a second mode control design using 2-D roughness, the roughness height of 50% boundary layer thickness, width of 2 boundary layer thicknesss and spacing of 10 have been chosen to optimize the damping mechanism.