Hypersonic Flow Research Articles

Three-Dimensional Shock Topology Detection Method via Tomographic Reconstruction

Shock waves and shock-shock interaction are typical phenomena in supersonic or hypersonic flows that have significant impacts on aerodynamic performance. To obtain a comprehensive understanding of the mechanism of shock wave interaction, shock wave detection (SWD) methods are required. However, it is often challenging for most current SWD methods to identify the relationship between shock waves (also known as shock topology). To address this issue, this paper proposes a novel three-dimensional shock topology detection method based on the tomographic reconstruction strategy. This method involves extracting parallel slices from the flow field, then utilizing a two-dimensional shock topology recognition algorithm to obtain shock lines. Shock bands are obtained by connecting shock lines for every two adjacent slices, and shock surfaces are generated by assembling shock bands. Interaction lines are also formed by connecting interaction points. The detected shock wave is a structure composed of “point-line-band-surface”, and the topology relationship with other shock waves is obvious. Numerical results show that the shock waves detected by the proposed method can be categorized into families. Moreover, the shock surfaces generated by this method are free of gaps, holes, and un-physical fragments, which is an improvement over existing SWD methods.

Modal Instabilities over Blunted Cones at Angle of Attack in Hypersonic Flow

The effects of nose radius and angle of attack on the linear modal amplification over blunt circular cones at hypersonic speeds are computationally investigated. The three-dimensional laminar flow solutions over a 1.5-m-long, 7°-half-angle cone with 5.080, 9.525, and 25.40 mm nosetip radii are computed for selected angle-of-attack values and freestream conditions that match the Mach 10 experiments at a freestream unit Reynolds number of 17.1 million per meter conducted within the Hypervelocity Wind Tunnel 9 at the Air Force Arnold Engineering Development Complex (AEDC). Results indicate that the linear amplification of stationary and traveling crossflow waves along inflection lines increases with the angle of attack and decreases with the nosetip radius. The trend in transition front with respect to increasing angle of attack and bluntness is found to be consistent with the predicted increase in the amplification factors for Mack’s second mode (MM) disturbances along the streamline trajectories for the small and medium bluntness cases. The increase in the MM amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray increases with the angle of attack and decreases with the nosetip radius. Furthermore, the transition -factors along the leeward ray are below 2 and, therefore, rather small to lead to transition.

Velocity Analysis of Supersonic Jet Flow in Double-Cone Ignition Scheme

Abstract In the double-cone ignition schemes (DCIS), the deuterium-tritium target shell is ablated and compressed by a high-power nanosecond laser in u-cones to generate plasmas. Under the actions of spherical symmetric compression and acceleration along the Au cone, they will be ejected out of the cone mouth and collide with each other. The plasmas witness conversion from kinetic energy to internal energy at the vertex of the geometric center of two Au cones that are symmetric to each other, because of which, fusion high-density plasmas are preheated. This key physical process has undergone experimental verification on the Shenguang-II upgraded facility of China. Apparently, the improvement and optimization of the velocity of plasmas in hypersonic jet flow at the cone mouth are crucial for the success of DCIS. In the DCI-R7 experiment for the Shenguang-II upgraded facility, a velocity yield of approximately 130-260 km/s was achieved for the plasmas at the cone mouth, with a result of nearly 300 km/s based on numerical simulation. In this paper, theoretical analysis is performed as regards the process, in which target shells are ablated and compressed by laser to generate high-velocity plasmas ejected through jet flow. Based on this analysis, the formula for the velocity of plasmas in supersonic jet flow at the cone mouth is proposed. This study also provides measures that are more effective for improving the kinetic energy of plasmas and optimizing energy conversion efficiency, which can serve as theoretical references for the adjustment and optimization of processes in subsequent experiments.

The Post-Shock Nonequilibrium Relaxation in a Hypersonic Plasma Flow Involving Reflection off a Thermal Discontinuity

The evolution in the post-shock nonequilibrium relaxation in a hypersonic plasma flow was investigated during a shock’s reflection off a thermal discontinuity. It was found that within a transitional period, the relaxation zone parameters past both the reflected and transmitted waves evolve differently compared to that in the incident wave. In a numerical example for the non-dissociating N2 gas heated to 5000 K/10,000 K across the interface and M = 3.5, the relaxation time determined for the transmitted wave is up to 50% shorter and the relaxation depth for both waves is significantly reduced, thus resulting in a weakened wave structure. The results of the extension into larger values of heating strength and the shock Mach numbers are discussed. The findings can be useful in the areas of research involving strong shocks interacting with optical discharges or other heated media on the scale where the shock structure becomes important.

A quasi-one-dimensional model for the stagnation streamline in hypersonic magnetohydrodynamic flows

The flow near the stagnation streamline of a blunt body is often attracted and analyzed by using the approximation of local similarity, which reduces the equations of motion to a system of ordinary differential equations. To efficiently calculate the stagnation-streamline parameters in hypersonic magnetohydrodynamic (MHD) flows, an improved quasi-one-dimensional model for MHD flows is developed in the present paper. The Lorentz force is first incorporated into the original dimensionally reduced Navier–Stokes equations to compensate for its effect. Detailed comparisons about the shock standoff distance and the stagnation point heat flux are conducted with the two-dimensional Navier–Stokes calculations for flows around the orbital reentry experiment model, including gas flows in thermochemical nonequilibrium under different magnetic field strengths. Results show that the shock curvature should be considered in the quasi-one-dimensional model to prevent accuracy reduction due to the deviation from the local similarity assumption, particularly for hypersonic MHD flows, where the shock standoff distance will increase with larger magnetic strength. Then, the shock curvature parameter is introduced to compensate for the shock curvature effect. A good agreement between the improved quasi-one-dimensional and the two-dimensional full-field simulations is achieved, indicating that the proposed model enables an efficient and reliable evaluation of stagnation-streamline quantities under hypersonic MHD flows.

Effect of surface radiation on the surface heat flux of two-dimensional hypersonic panel flow

This article investigates the effect of surface radiation on the surface heat flux of a two-dimensional hypersonic panel flow by comparing the surface heat fluxes of penetrating-radiative-equilibrium cases with those of constant-wall-temperature cases. A Du Fort–Frankel-type difference scheme is applied to cases with different external flow properties and is verified by comparing the results under a constant-wall-temperature boundary condition with self-similar solutions. Both laminar and turbulent flows are considered, and turbulence is modeled using a Baldwin–Lomax turbulence model with the assumption that full-scale turbulence is reached at the leading edge. The results show that the surface heat fluxes for laminar penetrating-radiative-equilibrium cases are greater by the order of 10% than those for constant-wall-temperature cases with the same temperature value at the corresponding points. Though turbulent instances are substantially more difficult, surface heat fluxes of penetrating-radiative-equilibrium cases are fairly similar to those of constant-wall-temperature cases. These results serve as the foundation for a brief discussion of how parameters affect the outcome and the proposal of a modified reference enthalpy method that can be used to predict heat flux when surface radiation causes a streamwise variation in wall temperature.

Heat transfer to proximal cylinders in hypersonic flow

Uncontrolled atmospheric entry of meteors, satellites, and spacecraft components often leads to their partial or complete demise. In this destructive process, driven by hypersonic ablation, reentry objects fragment, interact, and alter each other's aerothermal environment. The effect of these interactions on the heat transfer to the fragments has not been investigated, despite the heat transfer's importance in hypersonic ablation and reentry demise. This study focuses on the numerical investigation of heat transfer to proximal circular cylinders in a thermochemically frozen flow and in two dimensions. First, binary body configurations at Mach numbers 2, 4, and 8 revealed that the heat load and peak heat transfer can be augmented for either or both proximal bodies by +20% to −90% of an isolated body. Second, different clusters of five proximal bodies showed that the heat load to any given body can range from +40% to −90% of an isolated body. Moreover, the average heat load in a cluster is found to vary between +20% and −60% of an isolated body. Intuitively, clusters which are thin in the direction perpendicular to free-stream velocity and long in the direction parallel to the free-stream velocity have their cluster-averaged heat load reduced. In contrast, thick and thin clusters, in directions perpendicular and parallel to the free-stream velocity, are subject to an increased cluster-averaged heat load.

Flow alteration technique to mitigate wave drag by hemispherical aero-spike spike for large angle blunted cone in hypersonic flow

Flow alteration drag reduction by hemispherical aero-spike for a larger angle blunt cone at hypersonic speed (Mach = 8.0) is investigated by using ANSYS Fluent. Numerical studies were carried out using finite volume (FV) based compressible flow solver to solve the two-dimensional Reynolds-averaged Navier-Stokes (RANS) equation. In this study, the k-ωtwo-equation turbulence model was used. Air has been taken as flowing fluid for analysis. In this analysis, aero-spike tip of hemispherical shape has been taken. The length of the aero-spike spike is taken on the basis of four different lengths to base diameter (L/D) ratios, in which these ratios are 0.2, 0.5, 0.7 and 1.0 respectively. The shortest length aero-spike is 14 mm and the longest length is 105 mm. The numerical investigation found a drag reduction of about 23–76 % using aero-spike in a full planner 600 apex angle blunt cone.

Direct numerical simulation of hypersonic wall-bounded turbulent flows: An improved inflow boundary condition and applications

In this paper, the method of generating inflow turbulence based on turbulence fluctuation library (TFL) in direct numerical simulation (DNS) of the hypersonic turbulent boundary layer (TBL) is investigated. The application of the TFL method to the DNS of a supersonic TBL shows that, although there are significant differences in freestream between the TFL and the target TBL, the flow could successfully develop to the target TBL downstream as the fluctuations of TFL are suitably scaled and added to the DNS inflow. However, there is a “transition”-like recovery process from the inflow to the target turbulence. To deal with the defects of the thermodynamic fluctuations scaling laws in the current TFL method under the hypersonic TBL, new thermodynamic fluctuations scaling laws are theoretically derived by introducing the generalized Reynolds analogy. The application in the DNS of Mach 7.25 TBL shows that the new scaling laws for thermodynamic fluctuations are more rational and accurate than the previous ones. Furthermore, the study on the recovery process shows that the matching degree between the TFL and the target TBL on the friction Reynolds number (Reτ) is the dominant factor in determining the length of recovery distance. Guaranteeing the similar Reτ of the TFL and the target TBL can make the two possess similar coherence structures, which can avoid the distortion of the coherence structures at the inflow after spanwise and normal interpolation, prevent the process of Reynolds stress decay and readjustment downstream the inflow, and finally effectively shorten the recovery distance.

Active transition control by synthetic jets in a hypersonic boundary layer

We investigate by direct numerical simulation the active control of laminar-turbulent transition in a hypersonic flat-plate boundary layer at a freestream Mach number of 5.86. The control mechanism is a synthetic jet. Based upon the linear stability theory of Mack, in hypersonic flow the important path to transition involves a high-frequency, second-mode fundamental resonance. Through systematic investigation, we reveal that the forcing the boundary layer with a synthetic jet at appropriate combinations of amplitude and frequency suppresses the second mode and delays transition. To gain physical insights into the major control mechanism, we employ the momentum potential theory (MPT) to analyze the flows with and without control. Essentially, the underlying control mechanism relies on an intriguing effect of the synthetic jet via generating the outward radiated wave structures, which are identified to split the upstream acoustic and vortical components. The splitting treatment presents the second-mode energy to drop sharply after the flow passes through the synthetic jet slot. The MPT source-term analysis reveals that the significantly suppressed near-wall source terms are responsible for suppressing the second mode downstream. Compared with the vortical and thermal source terms, the acoustic source term is found to be suppressed most. The kinetic budget analysis further reveals that the splitting treatment is related to the non-parallel effect and the nonlinear interaction.

Roughness effect on hypersonic second mode instability and transition on a cone

Recent studies by Fong et al. [“Finite roughness effect on modal growth of a hypersonic boundary layer,” AIAA Paper No. 2012-1086, 2012; “Stabilization of hypersonic boundary layer waves using 2-D surface roughness,” AIAA Paper No. 2013-2985, 2013; “Numerical simulation of roughness effect on the stability of a hypersonic boundary layer,” Comput. Fluids 96, 350–367 (2014); and “Second mode suppression in hypersonic boundary layer by roughness: Design and experiments,” AIAA J. 53, 1–6 (2015)] have shown that finite roughness can attenuate Mack's second mode instability when placed at the discrete mode synchronization location for two-dimensional planar flow over a flat plate. However, more practical hypersonic flows are non-planer conical flows, and the roughness effect phenomenon in conical flows has not been extensively investigated. For that reason, this investigation research the ability of finite roughness strips to attenuate the second mode instability on a Mach 8 straight blunt cone with a freestream unit Reynolds number of 9 585 000. Two roughness configurations are studied: a single roughness strip and an array of six sequential strips. N-factor calculations determine the second mode frequency most likely to lead to turbulent transition, and the linear stability theory is used to determine the mode's synchronization location. In the unsteady simulations of the roughness configurations, a blowing-suction actuator introduces an upstream broadband Gaussian pulse. Fourier decomposition of the pulse's history shows that the single roughness strip attenuates frequencies higher than 208 kHz while lower frequencies are amplified. Likewise, the roughness array exhibits similar results, attenuating frequencies higher than 164 kHz and amplifying lower frequencies downstream. The results show that both configurations can delay second mode instability growth on a hypersonic blunt cone and possibly delay turbulent transition. However, investigations of the roughness effect's behavior downstream of the roughness configurations show that disturbance growth resumes and becomes more destabilizing to the boundary layer.

Influence of upstream wall slope on aerodynamic properties of a rectangular cavity in rarefied hypersonic flows

A comprehensive numerical study is performed to reveal influence of the upstream wall slope on aerodynamic properties as rarefied hypersonic flows pass through the two- and three-dimensional cavity configurations using the direct simulation Monte Carlo (DSMC) method. In this work, slope effect is obtained by the cavity upstream wall sloping upward and downward as well as bulging partially. Besides, cases of freestream with different velocities and altitudes are taken into account to examine the slope effect. It is found that cavity flow is transformed from the open type to the transitional and finally to the closed as the cavity upstream wall gradually slopes upward, promoting the heat energy exchange between gas molecules with the cavity floor, and meanwhile achieving a considerable rate of decrease of the peak heat flux on the reattachment corner. However, the upstream wall sloping downward or bulging partially seems to only have a slight effect on cavity flow and wall heat flux. It's worth noting that slope effect is closely related to velocity and altitude of the freestream, such as the greater the velocity and altitude, the stronger the slope effect. For the three-dimensional cavity configuration, influence of the upstream wall slope on cavity flow and wall heat flux is weaken considerably due to the three-dimensional effect. From the perspective of optimal design for the hypersonic thermal protection system, the cavity with an upward upstream wall is proposed because it not only reduces peak heat flux on the reattachment corner, but also decreases gas temperature inside the cavity.

Convergence rate of hypersonic similarity for steady potential flows over two-dimensional Lipschitz wedge

This paper is devoted to establishing the convergence rate of the hypersonic similarity for the inviscid steady irrotational Euler flow over a two-dimensional Lipschitz slender wedge in $$BV\cap L^1$$ space. The rate we established is the same as the one predicted by Newtonian-Busemann law (see (3.29) in [2, p 67] for more details) as the incoming Mach number $$\text {M}_{\infty }\rightarrow \infty $$ for a fixed hypersonic similarity parameter K. The hypersonic similarity, which is also called the Mach-number independence principle, is equivalent to the following Van Dyke’s similarity theory: For a given hypersonic similarity parameter K, when the Mach number of the flow is sufficiently large, the governing equations after the scaling are approximated by a simpler equation, that is called the hypersonic small-disturbance equation. To achieve the convergence rate, we approximate the curved boundary by piecewisely straight lines and find a new Lipschitz continuous map $$\mathcal {P}_{h}$$ such that the trajectory can be obtained by piecing together the Riemann solutions near the approximated boundary. Next, we derive the $$L^1$$ difference estimates between the approximate solutions $$U^{(\tau )}_{h,\nu }(x,\cdot )$$ to the initial-boundary value problem for the scaled equations and the trajectories $$\mathcal {P}_{h}(x,0)(U^{\nu }_{0})$$ by piecing together all the Riemann solvers. Then, by the uniqueness and the compactness of $$\mathcal {P}_{h}$$ and $$U^{(\tau )}_{h,\nu }$$ , we can further establish the $$L^1$$ estimates of order $$\tau ^2$$ between the solutions to the initial-boundary value problem for the scaled equations and the solutions to the initial-boundary value problem for the hypersonic small-disturbance equations, if the total variations of the initial data and the tangential derivative of the boundary are sufficiently small. Based on it, we can further establish a better convergence rate by considering the hypersonic flow past a two-dimensional Lipschitz slender wing and show that for the length of the wing with the effect scale order $$O(\tau ^{-1})$$ , that is, the $$L^1$$ convergence rate between the two solutions is of order $$O(\tau ^{\frac{3}{2}})$$ under the assumption that the initial perturbation has compact support.

Rotated-hybrid Riemann solver for all-speed flows

The contribution deals with numerical solution of compressible flows modeled by the Euler or the Navier–Stokes equations. Addressed flow regimes range from very low Mach number flows to hypersonic flows. The numerical solution is obtained by the finite volume method based on the rotated-hybrid Riemann solver method, where two different Riemann solvers are used in different directions. The combination of the HLL and the HLLC schemes leads to carbuncle-free numerical method. Both the HLL and the HLLC schemes are further modified for improved accuracy and stability in cases with very low Mach number flows. Second order accuracy in space is achieved by the piecewise linear WENO reconstruction while accuracy in time is maintained by the explicit two-stage TVD Runge–Kutta method. The proposed method is validated on many test cases including solution of subsonic flow over an airfoil with M=0.01 or hypersonic flow past circular cylinder with M=20.

The elastic scattering cross sections of the N[formula omitted]-H and N-H[formula omitted] collisions for the transport theory

The transport cross sections and integrated cross sections are calculated for interactions of hydrogen and nitrogen atoms and ions (in the ground and excited states) for use in plasma science and hypersonic flows theory. The data for the transport cross sections are given in the energy range typically needed (or broader) for those applications: 10−6-37 hartree (around 2.7⋅10−5eV-1 keV). The reported cross sections are calculated from the high quality ab initio potential energy curves with additional extrapolation for highly repulsive region. The methods and energy range are discussed in the context of LXCat and IAEA (ALADDIN and CollisionDB) databases. The uncertainty of data is also estimated.

Laser Applications to Chemical, Security, and Environmental Analysis: introduction to the feature issue.

The eighteenth topical meeting on Laser Applications to Chemical, Security, and Environmental Analysis (LACSEA) was held in Vancouver, Canada from 11-15 July 2022, as part of the Optica Optical Sensors and Sensing Congress in a hybrid format allowing on-site and online attendance. The meeting featured a broad range of distinguished papers focusing on recent advances in laser and optical spectroscopy. A total of 52 contributed and invited papers were presented during the meeting, including topics such as photo-acoustic spectroscopy, imaging, non-linear technologies, frequency combs, remote sensing, environmental monitoring, aerosols, combustion diagnostics, hypersonic flow diagnostics, nuclear diagnostics, fs/ps applications, and machine learning and computational sensing.

Artificial neural network-based streamline tracing strategy applied to hypersonic waverider design

Streamline tracing in hypersonic flows is essential for designing a high-performance waverider and intake. Conventionally, the streamline equations are solved after obtaining the velocity field over a basic flow field from simplified flow differential equations or three-dimensional computational fluid dynamics. The hypersonic waverider shape is generated by repeatedly applying the streamline tracing approach along several planes. This approach is computationally expensive for iterative waverider optimization. We provide a novel strategy where an Artificial Neural Network (ANN) is trained to directly predict the streamlines without solving the differential equations. We consider the standard simple cone-derived waverider using Taylor–Maccoll equations for the conical flow field as a template for the study. First, the streamlines from the shock are solved for a wide range of cone angle and Mach number conditions resulting in an extensive database. The streamlines are parameterized by a third-order polynomial, and an ANN is trained to predict the coefficients of the polynomial for arbitrary inputs of Mach number, cone angle, and streamline originating locations. We apply this strategy to design a cone-derived waverider and compare the geometry obtained with the standard conical waverider design method and the simplified waverider design method. The ANN technique is highly accurate, with a difference of 0.68% from the standard method in the coordinates of the waverider. The performance of the three waveriders is compared using Reynolds averaged Navier–Stokes simulations. The ANN-derived waverider does not indicate severe flow spillage at the leading edge. The new ANN-based approach is 20 times faster than the standard method.

Stability of hypersonic flow over a curved compression ramp

In this work, the stability of hypersonic flow over a curved compression ramp is studied using several stability analysis tools and direct numerical simulations (DNS). The free-stream Mach number and the unit Reynolds number are 7.7 and $4.2 \times 10^6$ m $^{-1}$ , respectively. Corner rounding is considered to alter the separation bubble flow so as to suppress the intrinsic instability of the compression-ramp flow. The variation of intrinsic instability is confirmed by global stability analysis. Subsequently, resolvent analysis is employed to examine the response of intrinsically stable flows to external disturbances. It is shown that the considered flows strongly amplify low-frequency streamwise streaks with a preferential spanwise wavelength. This result is verified using DNS by introducing a random forcing upstream of the separation point. Furthermore, both resolvent analysis and DNS demonstrate that the separation bubble contributes little to the selection of the spanwise wavelength of streamwise streaks. The combined effects of convective and intrinsic instabilities are also explored using DNS. A better agreement with experimental data is achieved after introducing upstream disturbances in an inherently unstable flow.

Data-driven framework for input/output lookup tables reduction: Application to hypersonic flows in chemical nonequilibrium

In this paper, we present a novel model-agnostic machine learning technique to extract a reduced thermochemical model for reacting hypersonic flows simulation. A first simulation gathers all relevant thermodynamic states and the corresponding gas properties via a given model. The states are embedded in a low-dimensional space and clustered to identify regions with different levels of thermochemical (non)-equilibrium. Then, a surrogate surface from the reduced cluster-space to the output space is generated using radial-basis-function networks. The method is validated and benchmarked on a simulation of a hypersonic flat-plate boundary layer with finite-rate chemistry. The gas properties of the reactive air mixture are initially modeled using the open-source Mutation++ library. Substituting Mutation++ with the light-weight, machine-learned alternative improves the performance of the solver by 50% while maintaining overall accuracy.

Numerical Investigation of Mach 2.5 Axisymmetric Turbulent Shock Wave Boundary Layer Interactions

Shock wave boundary layer interactions are common to both supersonic and hypersonic inlet flows. Wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 axisymmetric shock wave boundary layer interaction experiment at Glenn Research Center were carried out. A conical shock wave was generated with axisymmetric centerbodies with 16 deg half-angle cone. The centerbody radii were 9.2% and 14.7% of the test section diameter. The conical shock wave interacted with the turbulent boundary layer on the inside of the cylindrical test section. The experimental Reynolds number based on diameter was six million. For the simulations, the Reynolds number was reduced by a factor of 10 to lower the computational expense. The turbulent boundary layer separates for both centerbody radii and the separation is stronger for the larger centerbody radius. Frequency spectra of the spanwise-averaged wall-pressure coefficient reveal low-frequency content at Strouhal numbers based on separation length between 0.02 and 0.05 in the vicinity of the separation shock and mid-frequency content between 0.1 and 0.2 downstream of separation. A proper orthogonal decomposition captures spanwise coherent structures with a Strouhal number of 0.03–0.04 over the interaction region and streamwise coherent structures inside and downstream of the interaction with a Strouhal number of 0.1–0.4.