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Purpose Pokemon, also known as ZBTB7A, belongs to the POZ and Kruppel (POK) family of transcription repressors and is implicated in tumor progression as a key proto-oncogene. This present study aimed at determining the mechanism by which Pokemon inhibits transforming growth factor β (TGFβ)-Smad4 pathway-dependent proliferation arrest of breast cancer cells via specificity protein 1 (SP1). Methods Over-expressing plasmid or small interfering RNA (siRNA) transfection was used to regulate Pokemon levels. The EdU incorporation assay, MTS assay, and clone formation were used to identify the inhibitory effect of Pokemon siRNA on cell proliferation. Quantitative real-time polymerase chain reaction assay confirmed that Pokemon deletion inhibited the expression of proliferation-associated genes. The dual-luciferase reporter assay, electrophoretic mobility shift assay, and co-immunoprecipitation assay were used to analyze binding between Pokemon, Smad4, and SP1. Results Pokemon deletion induced proliferation arrest of breast cancer cells and inhibited the expression of proliferation-associated genes, especially Smad4. Pokemon bound with SP1 to interdict Smad4 promoter activity. Information on clinical samples was obtained from The Cancer Genome Atlas data, in which the Pokemon mRNA levels showed a negative correlation with Smad4 levels in different subtypes of breast cancer in two independent datasets. Conclusion We demonstrated that Pokemon binds to SP1 to down-regulate Smad4 expression, thereby promoting proliferation of breast cancer cells. This suggests that Pokemon is a potential TGFβ-signaling participant in breast cancer progression.

Designing a photoelectrochemical (PEC) immunosensor that can produce dual photocurrent signals which can refer to each other is a great importance but a big challenge. In this manuscript, a novel dual photocurrent signals immunosensor was constructed for the detection of α-fetoprotein (AFP). Unlike the usual method of using two composite materials to provide cathode and anode photocurrent respectively, this work applies only one compound of MIL-101 (Cr) and CdSe quantum dots (QDs). Thereinto, we found that the photocurrent polarity of MIL-101(Cr) would switch by adjusting applied voltage. And then CdSe QDs was introduced by simple ultrasound mixing to boost the dual photocurrent signals. Furthermore, in the composite of M&C, the electron transfer path between MIL-101(Cr) and CdSe QDs may switch between “Z-type” and “Ⅱ-type” by adjusting voltage. Benefiting by the dual signals, the proposed sensor can not only perform sensitively quantitative detection of α-fetoprotein (AFP), but also can intuitively estimate the accuracy and reliability of the test result by determining whether the corresponding relationship of “cathode photocurrent-analyte concentration-anode photocurrent” is established. The linear ranges of the sensing electrodes as cathode and anode are the same, both from 0.1 to 300 ng mL−1. The limit of detection (LOD) is 0.082 ng mL−1 (S/N = 3) when it used as an anode, and the LOD is 0.054 ng mL−1 (S/N = 3) when it served as cathode. Furthermore, this sensor showed acceptable stability, reproducibility, specificity, and feasibility of detecting AFP in human serum, which has broad development prospects in the early clinical diagnosis.

Patients with ovarian cancer are often in a hypercoagulable state and have a high risk of venous thrombosis, including deep vein thrombosis and pulmonary embolism. However, arterial thrombosis is relatively rare in ovarian cancer. We report a case a 46-year-old woman with ovarian clear cell carcinoma who developed arterial and venous thrombosis in the lower extremities as the first manifestation. Her arterial thrombosis-related ischemic symptoms were not responsive to anticoagulant treatment of low-molecular-weight heparin, but improved after neoadjuvant chemotherapy and surgery. Therefore, we hypothesize that the optimal therapy for arterial thrombosis in ovarian cancer is treatment for the underlying disease (i.e., ovarian cancer). A thorough investigation is required to determine the relationships between arterial thrombosis and ovarian cancer and antithrombotic treatments for ovarian cancer related-arterial thrombosis.

Background: Helicobacter pylori (Hp) is a unipolar, multi-flagellate, spiral,Gram-negative bacterium that is the only stomach bacterium found in humans.Hp is the leading cause of digestive tract disease and can cause stomach diseases, such as chronic gastritis, peptic ulcer, gastric cancer and gastric mucosa-associated lymphoma. Clinical studies have shown that Hp infection rates are significantly higher in patients with hyperlipidaemia than they are in non-hyperlipidaemic patients. In addition,many studies have shown that the serum vitamin D3 levelsare reduced in patients with obesity and hyperlipidaemia.Lower serum vitamin D3 levels diminish the role of fat cells in storing fat;as a result,fat is released into the bloodstream,leading to elevated blood lipid levels.However, a specificmechanism has not been proposed to explain how VD3 can reduce HP infection in patients with hyperlipidaemia by lowering blood lipids. Methods real-time PCR was used to detect the expression of gastrin in VDR, insign-2 and gastric tissues of mice in each group. The expression of jak, stat 3, cox2 and lkba protein in gastric tissue of mice was detected by western blotting. We also detected the expression of IL-6 and IL-8 in the gastric tissue of mice by Elisa method. The liver tissues of each group were stained with HE to observe the balloon-like changes of liver tissue, and the gastric tissues of mice in each group were stained by HE. The degree of gastritis in each group of mice. Conclusions In this article, we conducted animal experiments using VD3 as an intervention in hp-infected mice and then examined molecular changes in the VDR/Insig-2/SREBP-2 pathways in the mouse liver, the serum cholesterol levels, and the JAK-STAT3 and NF-ĸB pathways.Vitamin D3 further reduced the gastric mucosa cholesterol level,destroyed lipid rafts,and reduced assembly of IL-6 receptor subunits by lowering the circulating cholesterol levels, thereby blocking the JAK-STAT3 signalling pathway. Infection of the gastric mucosa by CagA-positive Helicobacter pylori reduced gastrin secretion and inhibited COX2 expression. Moreover, NF-κB activation and inflammatory cytokine expression were inhibited by blocking the JAK-STAT3 signalling pathway. Furthermore, the degree of infection of the gastric mucosal tissue was reduced.

A computational fluid dynamics study is conducted in which acoustic-like waves are observed emanating from the boundary layer of a Mach 8 slender blunt cone with a relatively low freestream enthalpy and a warm wall. The acoustic-like wave emissions are qualitatively similar to those attributed to the supersonic mode. However, the supersonic mode responsible for such emissions is often found in high-enthalpy flows with highly cooled walls, making its appearance here unexpected. Linear stability analysis on the steady-state solution reveals an unstable mode S (Mack's second mode) with a subsonic phase velocity and a stable mode F whose mode F- branch takes on a supersonic phase velocity. It is thought that the stable supersonic mode F- is responsible for the acoustic-like wave emissions. Unsteady simulations are carried out using blowing-suction actuators at two different surface locations. The analysis of the temporal data and spectral data using Fourier decomposition reveals constructive/destructive interference occurring between a primary wave packet and a satellite wave packet in the vicinity of the acoustic-like wave emissions. The constructive/destructive interference between the wave packets also appears to have a damping effect on individual frequency growth in both unsteady simulations. Based on this study's results and analysis, it is concluded that a supersonic discrete mode is not limited to high-enthalpy, cold wall flows and that it does appear in low-enthalpy, warm-wall flows; however, the mode is stable.

There has been little research into surface-ablation effects on hypersonic boundary-layer instability, and the current understanding of real-gas effects on hypersonic boundary-layer instability still contains uncertainties. The objective of the current work was to analyze the hypersonic boundary-layer transition process using linear-stability theory, in which surface ablation, as well as real-gas effects, is present, and the second mode is the dominant instability mode. A thermochemical nonequilibrium linear-stability-theory code with a gas-phase model that includes multiple carbon species, as well as a linearized surface graphite-pseudoablation model, is developed and validated. It is validated with previously published linear-stability analysis and direct-numerical-simulation results. A high-order method for discretizing the linear-stability equations is used, which can easily include high-order boundary conditions. The developed linear-stability code, with mean-flow solutions produced from a high-order...

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.

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.

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.

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.

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.

The numerical simulation of transitional and turbulent e ows in supersonic boundary layers often involves a physical process of a shock ‐disturbance wave interaction in complex multidimensional e owe elds. For such simulations, it is required that there be a high order of accuracy in capturing the shock waves without spurious numerical disturbances. Evaluation of the numerical oscillations generated behind a stationary bow shock by using high-order shock-capturing schemes in computing multidimensional steady supersonic e ow over a circular cylinderiscarriedout.Thenumerical methodsthatarestudied aretheTotalVariation Diminishing schemeand the Essentially Non-Oscillatory scheme. Although the general aerodynamic properties are appropriately captured by the shock-capturing schemes, it is shown that there are numerical oscillations in the gradients of the aerodynamic propertiesin the steady e owe eld behind the bow shock, such asfor vorticity. Thesespurious numerical oscillations in the e owe eld solution may hinder any attempt at tracking the propagation of physical disturbances behind the shock if unsteady simulations are carried out. They can be signie cant enough to pollute a e owe eld containing small physical disturbances. It is shown that the effects of grid ree nement do not reduce the oscillations but rather decreasetheir wavelength. It is also shown that, by roughly aligning the shock with the grid, the amplitude of these spurious oscillations can be reduced but not eliminated.

This paper presents the numerical simulations of transient flow under the interaction between freestream hotspot perturbations and bow-shock on Purdue’s blunt compression cone with the effects of different nosebluntness and various freestream Mach numbers. The basic flow conditions are referring to Boeing/AFOSR Mach-6 Quiet tunnel (BAM6QT) at Purdue University. The simulations in this paper consist of two cases of different nose-radii and two cases of different freestream Mach numbers. The detailed analysis of the receptivity mechanisms is carried out with linear stability theory (LST). The simulation results agree well with LST analyses. The parametric study is performed by the comparisons of the simulation results with the Mach-6 results that have been carried out previously. The study shows the sharper cone destabilizes the boundary layer, and the higher freestream Mach number stabilizes it. The receptivity mechanisms under both effects are shown to be the same: the second mode instability is induced by the fast acoustic waves behind the shock in the nose region.

In the past decade, passive control of hypersonic boundary-layer transition using porous coatings has been studied by theoretical analyses, experiments, and numerical simulations. It was found that porous coating stabilizes Mack’s second mode and destabilizes Mack’s first mode. However, there is only quite limited studies on the stabilization efficiency of porous coating. And no work has been reported about the thermochemical non-equilibrium effects of hypersonic flows on the stabilization efficiency. In this paper, we conduct numerical simulations on the passive control of hypersonic boundary-layer transition using regular porous coating. The stabilization of a Mach 5.92 flat-plate boundary layer is first studied for perfect gas flow. The results show that, at approximately the same porosity, regular coating is weaker in first-mode destabilization and second-mode stabilization than felt-metal coating. The porosity decrease of regular coating leads to even weaker first-mode destabilization and second-mode stabilization. The results also show that the first-mode destabilization weakens as the phase angle of admittance decreases and the thermochemical nonequilibrium of hypersonic flows may affect the stabilization efficiency of regular coating. Therefore, numerical simulations based on perfect gas flow may not be enough. The effects of thermochemical non-equilibrium flow including internal energy excitations, translation-vibration energy relaxation, and chemical reactions among different species need to be considered. We have developed a new high-order shock-fitting solver for non-equilibrium flow simulations based on the 5-species air chemistry and recently thermal non-equilibrium models. The code package has been tested and is being applied to numerical simulation of a Mach 12.56 boundary layer over a blunted wedge of a half angle 20 degree. Thermochemical non-equilibrium effects of hypersonic flows on the steady base flow is investigated by comparing numerical results of perfect gas flows and thermochemical nonequilibrium flow. Unsteady simulations on the passive control of hypersonic non-equilibrium boundary-layer transition using regular porous coating are currently ongoing.

As a continued effort in investigation of hypersonic boundary layer transition induced by fresstream waves, direct numerical simulations are performed on blunt cones over Mach 5.5 and Mach 6 flows. In this paper, a new approach is introduced to simulate the hypersonic flow over blunt cone from laminar to the breakdown stage in transition. The disturbance waves are introduced into freestream instead of generating within the boundary layer. The new approach breaks the direct numerical simulation into three parts: mean flow computation, linear receptivity simulation and nonlinear breakdown simulation. With this approach, it is possible to directly link the freestream receptivity process to the final breakdown stage so that the complete transition process can be better understood. The ultimate goal is to develop a robust numerical tool to carry out hypersonic boundary layer flow simulation over blunt cone to the non-linear breakdown stage and use it for future transition studies.

disturbances into account and highly relies on the experimental correlation. Due to the rapidly growing large scale computation capability, direct numerical simulation (DNS) has become a promising tool in assisting us to unveil the mystery of hypersonic boundary layer transition. In this paper, a new approach is presented to simulate the hypersonic boundary layer flow over a cone from laminar to the nonlinear breakdown using disturbance induced from the freestream waves. The goal is to develop a robust numerical tool that using a built linear receptivity library as inflow forcing to simulate the hypersonic boundary later flow to the non-linear breakdown stage and study the effects of freestream disturbance levels to the location of transition.

Results are presented for direct numerical simulations of hypersonic flow over an isolated roughness element located on a flat plate. The simulations were performed on a Cartesian grid with the roughness geometry imposed by a third-order cut-cell method. The simulations also utilize a fifth-order WENO scheme to capture any steep gradients in the flow created by the roughness. The combination results in a globally fourth-order scheme. Two-dimensional and three-dimensional test cases show very good agreement with results obtained from a body-fitted grid simulation. Preliminary results for Mach 6 flow over a k/� = 0.75 cylindrical roughness are reported. In addition, results from a series of simulations of Mach 6 flow over a three-dimensional Gaussian roughness performed with a body-fitted grid are presented.

Passive control of hypersonic boundary-layer transition using feltmetal and regular porous coatings has recently motivated many researches. The stabilization effects of porous coating on hypersonic boundary layers over flat plates and cones has been demonstrated by theoretical analyses, experiments, and numerical simulations. Previous results showed that both feltmetal and regular porous coatings significantly stabilize Mack’s second mode, whereas Mack’s first mode is moderately destabilized by porous coatings. Up to now, there is no work reported on how to deal with the first mode destabilization and increase the stabilization efficiency of porous coating. In this paper, we analyzed the porous coating admittance and noticed it was the phase angle of admittance that induces the phase discrepancy between the velocity and pressure perturbations on the wall. The effect of the admittance phase angle on Mack’s first mode destabilization and Mack’s second mode stabilization is then studied by series of numerical simulations for artificial porous coatings. The results showed that the phase angle of porous coating admittance has a critical role on boundary-layer stabilization. Specifically, the destabilization of Mack’s first mode is closely related to the phase angle of porous coating admittance, i.e., the destabilization effect decreases with the phase angle of admittance decreasing. The relation between admittance phase angle and first mode destabilization is quite useful for the design of new porous coating.

The stabilization effect of porous coating on hypersonic boundary-layer flows has been studied by theoretical analyses, experiments, and numerical simulations. However, most of the previous work considered either felt-metal porous coating or regular porous coating. Comparison of the two types of porous coatings was only done in experiments. In this paper, mode S growth over felt-metal and regular porous coatings of a Mach 5.92 flow is obtained by numerical simulations to study the effect of different porous coatings. The results show that the two types of porous coatings both destabilize mode S in Mack’s first mode region. When mode S propagates into Mack’s second mode region, porous coatings initially stabilize it before destabilizing it further downstream. At approximately the same porosity, regular porous coating is weaker in first mode destabilization, initial second mode stabilization, and downstream second mode destabilization. In addition, the wall-normal velocity disturbance induced by regular porous coating is significantly weaker than that induced by felt-metal porous coating. For regular porous coating, decreasing of porosity leads to even weaker first mode destabilization, initial second mode stabilization, downstream second mode destabilization, and wall-normal velocity disturbance on the surface. Further studies are needed for the exact mechanism behind the complex effect of porous coating on mode S in Mack’s second mode region.

The stabilization effect of porous coating on hypersonic boundary layers over flat plates and cones has been studied by theoretical analyses, experiments, and numerical simulations. The results show that porous coating slightly destabilizes the Mack first mode whereas it significantly stabilizes the Mack second mode. In previous numerical simulations, either felt-metal porous coating or regular structure porous coating is considered. Comparison of the two types of porous coatings is only done in experiments. In this paper, the impact of porous surface on hypersonic boundary layer instability is studied, focusing on the effect of different porous coating. Stability simulations are carried out by introducing disturbances corresponding to mode S at a cross-section of the boundary layer near the leading edge. It is found that the two types of porous coatings both destabilize the Mack first mode whereas they stabilize the Mack second mode. For the specific cases considered in current paper, the overall effect of porous surface is destabilizing mode S. The peak amplitude of mode S propagating over porous surface is slightly higher than that of mode S propagating over solid surface. At approximately the same porosity, regular structure porous coating is weaker in first mode destabilization and second mode stabilization than felt-metal porous coating. Therefore, the two types of porous coatings lead to approximately the same peak amplitude of mode S. However, the wall-normal velocity disturbance on the wall induced by regular structure porous coating is significantly weaker than that induced by felt-metal porous coating. For regular structure porous coating, the results also show that decreasing of porosity leads to even weaker first mode destabilization, second mode stabilization, and wall-normal velocity disturbance on the wall.

Hypersonic boundary-layer transition can be affected significantly by surface roughness. Many important mechanisms which involve transition induced by arbitrary roughness are not well understood. Direct numerical simulation is broadly applied for investigating the roughness induced instability and transition in recent decade. But due to the complex computational geometry with embedded 3-D roughness, the use of body fitted curvilinear grids could prove to be very difficult. In our previous papers [44,45], we proposed a new high-order cut cell method to overcome the natural complexities in grid generation around curvilinear surface. The new algorithms of up to 4 () Oh accuracy have been derived and applied to simulate two-dimensional boundary-layer transition with finite surface roughness. Compared with two-dimensional roughness, the flow with three-dimensional isolated and distributed roughness is more complicated in analyzing and simulating. In this paper, we extend our previous new high order cut cell method to simulate three-dimensional hypersonic flow with finite surface roughness. The flow structure in three-dimensional boundary layer is investigated by using the new numerical approach.

The laminar-turbulent transition of hypersonic boundary layers has a significant effect on drag calculation and aerothermal analysis of hypersonic vehicles. Surface roughness has been shown to have profound effects on boundary layer transition. Recent research has shown that one possible explanation to roughness induced bypass transition is the transient growth theory. However, there has been very few direct numerical simulation studies on transient growth in hypersonic boundary layers. This paper presents some results in on-going numerical simulation study on the transient growth of a Mach 5.92 flat-plate boundary layer to small three-dimensional surface roughness. The main objective is to study the effects of nonparallel flow and spanwise wave on transient growth. The freestream flow conditions are the same as those used in our previous study [1] . The responses of the boundary layer to surface roughness are computed by solving three-dimensional Navier-Stokes equations with a fifth-order shockfitting method and a Fourier collocation method. The numerical results show that no transient growth is observed when surface roughness is located in parallel flow region, which indirectly demonstrates that nonparallel flow effects enhance transient growth. It is also shown that surface roughness with a smaller spanwise wave number is more efficient in inducing disturbances with respect to energy norm.

Direct numerical simulations on the receptivity of hypersonic boundary layers over a flat plate and a sharp wedge were carried out with two-dimensional periodic-in-time wall blowing-suction introduced into the flow 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 for the filtering out of the stable and unstable modes hidden behind perturbations of another physical nature. I. Introduction The progress being made in computational fluid dynamics provides an opportunity for the reliable simulation of such complex phenomena as laminar-turbulent transition. The dynamics of flow transition depends on the instability of small perturbations excited by external sources. Computational results provide complete information about the flow field that would be impossible to measure in real experiments. Recently, a method of normal mode decomposition was developed for two- and three-dimensional perturbations in compressible and incompressible boundary layers. 1–3 In Ref. 4, the method was applied to the theoretical analysis of the perturbation flow field in the vicinity of the blowing-suction actuator obtained from direct numerical simulation (DNS). The results demonstrated very good agreement between the amplitudes of the modes filtered out from the DNS data and those solved by linear theory of the flow receptivity to wall blowing-suction. In the present work, we apply the multimode decomposition to DNS results downstream from the blowingsuction actuator in hypersonic boundary layers past a flat plate and a sharp wedge to compare the amplitudes of the modes found from the computations with the prediction of linear stability theory. II. Outline of the multimode decomposition The method of multimode decomposition of perturbations having a prescribed frequency is based on the biorthogonal eigenfunction system for linearized Navier-Stokes equations. 3 For the clarity of further discussion, we reproduce the main definitions necessary for discussing the present work. We consider a compressible two-dimensional boundary layer in Cartesian coordinates, where x and z are the downstream and spanwise coordinates, respectively, and coordinate y corresponds to the distance from the wall. We write the linearized Navier-Stokes equations for a periodic-in-time perturbation (the frequency

The stabilization effect of porous coating on hypersonic boundary layers over flat plates and cones has been successfully demonstrated by theoretical analyses, experiments, and numerical simulations. It has been found that porous coating slightly destabilizes the Mack first mode whereas it significantly stabilizes the Mack second mode. In previous studies, porous coating covers either the entire flat plate or the surface around half the cone circumference. The effect of porous coating locations on boundary layer stabilizations has not been considered. Furthermore, the destabilization effect of porous coating on the Mack first mode has not been studied in detail. In this paper, the stabilization of a Mach 5.92 flat-plate boundary layer using felt-metal porous coating is studied, focusing on the effect of porous coating locations and the destabilization of the Mack first mode. Stability simulations are carried out by introducing disturbances corresponding to mode S at a cross-section of the boundary layer near the leading edge. In a series of simulations, porous coatings are put both upstream and downstream of the synchronization point. It is found that porous coating location has an important effect on the stabilization of mode S. Disturbances are destabilized when porous coating is located upstream of the synchronization point whereas they are stabilized when porous coating is downstream of the synchronization point. For felt-metal porous coating, destabilization of the Mack first mode is significant so that disturbances are slightly destabilized when porous coating are put on the whole flat plate. The results indicate that an efficient way to s hypersonic boundary layers is to put porous coating downstream of the synchronization point.

The objective of this research is to conduct DNS studies of hypersonic boundary layer receptivity, transient growth and transition with surface roughness. The main approach is to use DNS as a research tool to study the boundary layer receptivity and transient-growth mechanisms in hypersonic flows, including the development of numerical algorithms and parallel computer codes of higher order numerical methods for the simulation of hypersonic flows with surface roughness of finite heights. During the three-year period, we have conducted DNS studies on the hypersonic boundary layer flows over flat plates and blunt cones. A new high-order cut-cell method has been developed for the numerical simulation of hypersonic boundary layer transition with finite height surface roughness. The method has been applied to the numerical simulations of two-dimensional hypersonic flows over a flat plate. Furthermore, the stabilization effect of the surface porous coating over a flat plate is extensively studied by series of numerical simulations. We also collaborate with Prof. Tumin in the University of Arizona to compare numerical and theoretical results on receptivity of a Mach 5.92 flow over a flat plate to wall blowing-suction, and to analyze the nonparallel flow effect.

The effect of porous coating on the instability of a Mach 5.92 flat-plate boundary layer is investigated by a combination of numerical simulation and theoretical analysis. It is motivated by Fedorov et al.’s experimental and theoretical studies considering the effect of an ultrasonically absorptive coating on hypersonic boundary-layer stability [1] and our previous numerical simulations on hypersonic boundary-layer receptivity [2, 3] . The steady base flow is obtained by solving compressible Navier-Stokes equations with a combination of a fifth-order shock-fitting method and a second-order total variation diminishing (TVD) scheme. Stability simulations consist of two steps: 1. periodic disturbances corresponding to single boundary layer wave (mode F or mode S) are superimposed on steady base flow at a cross-section of the boundary layer to show spatial development of the wave; 2. porous coating, modelled by pressure oscillation related wall blowing-suction, is used downstream of the superimposed wave to investigate its effect on boundary-layer instability. Stability characteristics of the hypersonic boundary layer are analyzed by linear stability theory. The numerical results show that porous coating only has local effects on the spatial developments of mode S and mode F. In case of mode S, porous coating destabilizes the Mack first mode whereas it stabilizes the Mack second mode. All the six cases of stability simulations show a larger peak amplitude of mode S, which means destabilization of the Mack first mode is quite significant. In case of mode F, all three cases of stability simulations show that mode F changes to mode S near the synchronization point and mode F is generally stabilized in porous region

For high-speed flow over blunt bodies, gas temperature increases dramatically. As a result, gas molecules start dissociating, and vibration and electron energy modes are being excited. Gas becomes thermally nonequilibrium and chemically reactive. Effects of vibration and electron energy excitations, translation-vibration energy relaxation, and chemical reactions among different species need to be considered in numerical simulations. In current paper, a CFD solver based on a fifth-order WENO scheme is developed for the simulation of nonequilibrium and reactive high-speed flows, as part of the ADPDIS3D computer code package. The code is implemented based on a two-temperature model. It is assumed that translation and rotation energy modes are in equilibrium at the translation temperature whereas vibration and electron energy modes are in equilibrium at the vibration temperature. The code is validated and applied to numerical simulations of hypersonic nonequilibrium flows.

In this paper, effect of porous coating and its location on the spatial developments of mode S and mode F in a Mach 5.92 flat-plate boundary layer is investigated by linear stability theory (LST) and direct numerical simulation (DNS). The current study is motivated by Fedorov et al.’s experimental and theoretical studies of the effect of an ultrasonically absorptive coating on hypersonic boundary-layer stability [1] and our previous numerical simulations on hypersonic boundary-layer stabilities [2, 3] . The mean flow is calculated by solving compressible Navier-Stokes equations with a combination of a fifth-order shock-fitting method and a second-order total variation diminishing (TVD) scheme. The stability characteristics of the hypersonic boundary layer is analyzed by LST. Stability simulations based on mean flow consist of three steps: 1. disturbances corresponding to a specific boundary layer wave (mode S or mode F) at a frequency of 100 kHz are introduced at the inlet of the computational domain to study its spatial development; 2. porous coating is used to investigate its effect on the spatial development of mode S and mode F; 3. multiple porous coatings are used to study the effect of porous coating location on the spatial development of mode S and mode F. The same theoretical model of porous coating as that in Fedorov et al.’s paper is used, i.e., porous coating is modelled by pressure perturbation related wall blowing-suction. For all the cases considered in current paper, porous coatings are located upstream of the corresponding synchronization point of mode S and mode F at x ∗ = 0.33184 m. The numerical results show that porous coating and its location have significant effects on the spatial developments of mode S and mode F. No matter which mode of mode S and mode F is introduced at the inlet, porous coating increases pressure perturbation amplitude downstream of the porous region. With the number of porous coatings increasing, pressure perturbation amplitude keeps increasing. In case of mode S, the effect of porous coating on mode S in porous region is not consistent. With the location of porous coating shifting from upstream to downstream, the effect of porous coating on mode S changes from destabilization to stabilization. In case of mode F, all cases show that mode F changes to mode S near the synchronization point and mode F is stabilized in porous region.

The effects of surface roughness on the stability of hypersonic flow are of great importance to hypersonic vehicles. The overall goal of our research is to provide a better understanding of the effects of surface roughness on transitional and turbulent hypersonic boundary layers. Direct numerical simulations have been performed of flow in the Boeing/AFOSR Mach 6 quiet wind tunnel at Purdue University. Three-dimensional simulations were performed of the nozzle without any roughness and 2D simulations were performed with an isolated roughness element located on the wall of the nozzle. Two different roughness heights were simulated, both of which were less then the local undisturbed boundary layer thickness. The roughness elements produced a combination of Mach wave and expansion fan in the nozzle. The influence of the roughness elements on the flow is investigated in this paper. It is found that the presence of the elements has a substantial effect on the boundary layer and stability profiles downstream of the element.

*† Thermal protection layer has been commonly used to protect the reentry vehicles from the harsh conditions encountered at hypersonic velocities. Numerical simulations and experiments are used to predict the amount of this protective layer needed so that vehicles safely traverse the harsh zone. Prediction of ablation requires modeling of hypersonic flows coupled with reacting gas chemistry and the response of the solid protection layer to the extreme conditions. Current computational codes modeling chemistry with hypersonic flow are coupled with codes predicting the response of the protection layer. In addition there has been limited study on effects of chemical Nonequilibrium in flow the field and surface ablation on laminar-turbulent flow transition of hypersonic boundary layers. Such an integrated tool would facilitate the study of transition in chemically reacting ablative flow fields. Therefore it is our aim to develop an integrated high order code which solves the material response along with the flow with thermal and chemical non equilibrium with species ionization. In this paper we present results of shock capturing nonequilibrium solver with surface reactions for modeling ablation. The code has 11 species reaction model to account for carbon species.

The accurate prediction of laminar-turbulent transition in hypersonic boundary layers is a critical part of the aerodynamic heating analyses on hypersonic vehicles. Since 1950s, extensive wind-tunnel and flight-test experiments on boundary layer transition and a number of stability experiments have been conducted for hypersonic flows over circular cones [1, 2]. Schneider [1, 2] did an extensive review of the existing literature on these experiments for both flight test and wind tunnel experiments. Though many of these experiments did not measure transition mechanisms and were carried out in noisy wind tunnels, the experimental studies have led to better understanding of the effects on transition of many parameters, including nose bluntness, Mach number, freestream noise, surface and stagnation temperatures, freestream unit Reynolds numbers, cone half angles, angles of attack, and surface roughness, etc. Nevertheless, so far, the effects of many of these parameters on transition are still not well understood.

Understanding of transition of hypersonic boundary layers is critical for future development of re-entry vehicles. The receptivity of such a boundary layer to external flow disturbances is not well understood, particularly when the surface has finite-sized roughness. In this work, we explore the use of shock capturing for study of hypersonic boundary layer receptivity. In such problems, the interaction of the low-amplitude disturbances with shock structures may not be properly accounted for due to the localized dissipation inherent in the method. Previously-developed previously-developed WENO-based third- and fifth order shock capturing methods are applied to a receptivity problem, the Mach 4.5 flow past a flat plate with a sharp leading edge. This problem has been studied previously with a high-order shock fitting scheme. The shock-fitted solution is used to evaluate shock capturing schemes in terms of accuracy, eciency and stability. Also, the significance of the leading edge singularity will be studied, which was not possible with a shock fitting method.

A linear stability analysis is conducted to investigate hypersonic flows over blunt circular cones with a Mach 5.5 free-stream. The flow conditions are chosen to be the same as those of Stetson’s boundary layer transition experiments performed in 1967. The transition reversal phenomenon with respect to the nose bluntness was first observed by Stetson in these experiments. Therefore, the purpose of this study is to investigate the effects of nose bluntness on the linear stability of the boundary layer. For the case that transition reversal was observed experimentally, the second mode N factor is found to be too small to trigger the transition at the location reported in the experiment. The results show no bluntness induced reversal effects based on the second mode growth. In addition, another case adapted to Stetson’s Mach 8 experiment in 1984 is studied to investigate the second mode characteristics under different flow conditions.

According to a previous study on the receptivity of a hypersonic flat-plate boundary layer to surface roughness [1] , no very strong transient growth was induced by small amplitude surface roughness. Furthermore, the technique of linear boundary condition transfer was proved to be valid only when the height of surface roughness was smaller than approximately one thirty-fifth of the boundary layer thickness. In order to excite stronger transient growth, one can use steady blowing-suction disturbance or finite height roughness. In this paper, the receptivity of the same hypersonic flat-plate boundary layer to steady three-dimensional blowing-suction is studied by numerical simulations. The free-stream flow conditions are the same as those of Maslov et al.s leading-edge receptivity experiment [2] . Steady base flow is firstly computed by solving compressible Navier-Stokes equations with a combination of a fifth-order shock-fitting method and a second-order total variation diminishing (TVD) scheme. In receptivity simulations, steady blowing-suction periodic in spanwise direction is introduced on the flat plate. The subsequent responses of the hypersonic boundary layer are simulated by solving three-dimensional Navier-Stokes equations with a fifth-order shock-fitting method and a Fourier collocation method. Effect of blowing-suction amplitude on the receptivity is investigated by considering three cases of receptivity simulations. The numerical results show that counter rotating streamwise vortices are induced by steady blowing-suction. Compared with small amplitude surface roughness, steady blowing-suction is more efficient in i nducing streamwise vortices. Transient growth excited by steady blowing-suction is stronger than that excited by small amplitude surface roughness. As a byproduct, it is noticed that the linear properties of steady blowing-suction disturbance are valid up to the amplitude of ǫ = 1.0 × 10 −3 . For even higher amplitude steady blowing-suction, the nonlinear effect becomes important. New method is needed to measure the transient growth.

Thermal protection layer has been commonly used to protect the reentry vehicles from the harsh conditions encountered at hypersonic velocities. Numerical simulations and experiments are used to predict the amount of this protective layer needed so that vehicles safely traverse the harsh zone. Prediction of ablation requires modeling of hypersonic flows coupled with reacting gas chemistry and the response of the solid protection layer to the extreme conditions. Codes doing chemistry with supersonic flow are coupled with codes predicting the response of the protection layer. However it would be more useful to integrated codes which solve both the hypersonic flow and the response of the surface to such flows. Therefore it is our aim to develop an integrated code which solves the material response along with the flow with non equilibrium chemistry, vibrational excitation, species dissociation and ionization. For the present paper we present our preliminary results from a high order Navier Stokes solver which includes nonequilibrium reaction flow with two temperature vibrational energy model.

The objective of this research project is to compare our numerical simulation solutions with available experimental or theoretical results on hypersonic boundary layer receptivity and stability; and to conduct extensive DNS studies on the flow mechanisms of hypersonic boundary layer receptivity and stability. During the three-year period, we have conducted extensive DNS studies on the receptivity of hypersonic boundary layer flows over a sharp wedge, a flat plate, a blunt cone, and the FRESH aeroshell. DNS studies are compared with Stetson's 1984 stability experiment on Mach 7.99 flow over a blunt cone, and Maslov's leading-edge receptivity experiment on Mach 5.92 flow over a flat plate. Our numerical studies have been validated to be of high accuracy and led to further understanding of hypersonic boundary layer receptivity mechanism. Such understanding can lead to better tools for the prediction and control of high-speed boundary layer transition.

A high -order shock -fitting, finite -difference, direct - numerical -simulation (DNS) Navier -Stokes computer code has been under development. The code has b een developed specifically as a tool for investigating the attachment -line and crossflow boundary layer transition mechanisms associated with supersonic/ hypersonic airflows past swept/tapered wing geometries. This paper briefly introduces the importance o f studying the fundamental attachment -line and crossflow vortice swept -wing boundary layer instabilities mechanisms and discusses the reasons for approaching the problem using DNS simulations. Preliminary results from several DNS simulations of Mach 5.1 ai rflow about various parabolic cross -section wing leading edge geometries, differing in leading edge sweep angle and wing taper ratio, will be shown. Also to be shown are initial results from unsteady receptivity simulations where the mean flow solution is excited by a standing acoustic or vorticity wave introduced in the freestream ahead of the bow shock created in front of a swept parabolic leading edge in supersonic flow. GENERAL INTRODUCTION For the supersonic/ hypersonic air -vehicle designer being able to predict accurately using analytical or computational methods the laminar -turbulent boundary layer (B.L.) transition locations about practical 3 -D geometries is extremely desirable. Skin friction drag and surface heating rates are two important paramete rs which need to be predicted accurately, the first for range performance considerations and the second to insure vehicle safety by designing an appropriate thermal protection system capable of handling the intense surface heating, which yet will be light enough not to impact the vehicles usable payload very much. Both the skin friction and the surface heating rates are dependent on the type of boundary layer near the vehicles surface. A laminar boundary layer produces less skin friction and lower surface h eating rates than a turbulent boundary layer.

This paper presents a numerical simulation study of the receptivity to weak freestream acoustic waves for a Mach 5.941 axisymmetric flow over a flared cone with a 5' half-angle. The flow conditions and geometry are the same as those of the hypersonic boundary layer stability experiments carried out in NASA Lan ley Hypersonic Quiet Tunnel by Lachowicz et al. &? and transition experiments by Horvath et al. [51. In hypersonic boundary-layer flow over a blunt cone, the process of receptivity to freestream disturbances is altered considerably by the presence of the bow shock followed by an entropy layer. It is crucially important that the interaction of disturbance waves with the bow shock is accurately computed in numerical simulations. In the present study, both steady and unsteady flow solutions are obtained by computing the full Navier-Stokes equations with a fifth-order shock-fitting finite difference scheme, which is able to account for the effects of bow-shock/freestream-disturbance interaction accurately. Whenever possible, the current numerical results are compared with published experimental results. In addition, a normal-mode linear stability analysis is used to study the receptivity properties of the hypersonic boundary layer with the effects of the adverse pressure gradient and surface curvature. The focus of the study is on the generation of the second Mack modes in the boundary layer due to freestream acoustic disturbances, and the effects of surface curvatures of the flared cone on the receptivity process.