In the realm of dynamic separation problems, the motion of a body triggered by shock interactions is a common phenomenon. This is particularly important in terms of the safe separation of two-stage-to-orbit vehicles, where the motion must remain stable despite long-distance disturbances from shock waves. The flow field in these cases is complex, marked by interactions between hypersonic shock waves and a moving boundary. This leads to significant unsteady effects due to the body's translation and rotation over extended distances. Existing simulation techniques fall short in rapidly and accurately predicting the aerodynamic force and thermal properties for these problems, largely due to the overwhelming computational demands that result from oversize computational domains and the necessity of grid deformation. This paper presents a novel non-deforming grid method to address these challenges. The central concept is to anchor the reference frame to the moving object itself and to approach the problem from a non-inertial frame perspective. This accounts for the motion of the object solely via the inertial source term, circumventing the complexities of mesh manipulation typically required to link flow and motion equations. The moving shock boundary is designed to be closely compatible with selected shock-captured schemes, which reduces non-physical oscillations compared to the traditional method of direct assembly with theoretical shock relations. Other boundary conditions and the solution process are also refined to specifically target the unsteady, shock-dominated flow. These modifications significantly alleviate the computational burden. The effectiveness of the proposed method is demonstrated through several test cases. To showcase the method's practical application, a scenario is simulated wherein an ellipse is dislodged from a wedge by an incident shock wave, covering a long distance. These tests confirm the method's feasibility in aerospace engineering problems.

No AccessTechnical NotesExperimental Study on Aerodynamic Heating of Hypersonic Boundary-Layer BlowingZongxian Li, Meikuan Liu, Guilai Han, Dagao Wang and Zonglin JiangZongxian Li https://orcid.org/0000-0002-7372-2552Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Meikuan Liu https://orcid.org/0000-0002-2900-8861Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Guilai Han https://orcid.org/0000-0001-6442-2577Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Dagao WangChinese Academy of Sciences, 100190 Beijing, People's Republic of China and Zonglin JiangChinese Academy of Sciences, 100190 Beijing, People's Republic of ChinaPublished Online:3 Sep 2024https://doi.org/10.2514/1.J064454SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookXLinked InRedditEmail About References [1] Camillo G. P., Wagner A., Dittert C., Benjamin L., Wartemann V., Neumann J. and Hink R., "Experimental Investigation of the Effect of Transpiration Cooling on Second Mode Instabilities in a Hypersonic Boundary Layer," Experiments in Fluids, Vol. 61, No. 7, 2020, p. 162. https://doi.org/10.1007/s00348-020-02994-8 Google Scholar[2] Bogdanoff D. W., "Advanced Injection and Mixing Techniques for Scramjet Combustors," Journal of Propulsion and Power, Vol. 10, No. 2, 1994, pp. 183–190. https://doi.org/10.2514/3.23728 LinkGoogle Scholar[3] Yang H.-S., Liang H., Guo S.-G., Tang M.-X., Zhang C.-B., Wu Y. and Li Y.-H., "Research Progress of Hypersonic Boundary Layer Transition Control Experiments," Advances in Aerodynamics, Vol. 4, No. 1, 2022, p. 18. https://doi.org/10.1186/s42774-022-00105-1 Google Scholar[4] Miro F. M. and Pinna F., "Injection-Gas-Composition Effects on Hypersonic Boundary-Layer Transition," Journal of Fluid Mechanics, Vol. 890, May 2020, p. R4. https://doi.org/10.1017/jfm.2020.129 Google Scholar[5] Schneider S. P., "Hypersonic Boundary-Layer Transition with Ablation and Blowing," Journal of Spacecraft and Rockets, Vol. 47, No. 2, 2010, pp. 225–237. https://doi.org/10.2514/1.43926 LinkGoogle Scholar[6] Schmidt B. E., Bitter N. P., Hornung H. G. and Shepherd J. E., "Injection into Supersonic Boundary Layers," AIAA Journal, Vol. 54, No. 1, 2016, pp. 161–173. https://doi.org/10.2514/1.J054123 LinkGoogle Scholar[7] Fedorov A. V., Soudakov V. and Leyva I. A., "Stability Analysis of High-Speed Boundary-Layer Flow with Gas Injection," 7th AIAA Theoretical Fluid Mechanics Conference, AIAA Paper 2014-2498, 2014. LinkGoogle Scholar[8] Schmidt B. E. and Shepherd J. E., "Stability of Supersonic Flow with Injection," AIAA Journal, Vol. 57, No. 12, 2019, pp. 5230–5240. https://doi.org/10.2514/1.J058080 LinkGoogle Scholar[9] Ifti H. S., Hermann T., Ewenz Rocher M., Doherty L., Hambidge C., McGilvray M. and Vandeperre L., "Laminar Transpiration Cooling Experiments in Hypersonic Flow," Experiments in Fluids, Vol. 63, No. 6, 2022, p. 102. https://doi.org/10.1007/s00348-022-03446-1 CrossrefGoogle Scholar[10] Li Z., Liu M., Han G., Wang D. and Jiang Z., "Numerical Simulation of Hypersonic Flat-Plate Boundary-Layer Blowing Control," Physics of Fluids, Vol. 35, No. 12, 2023, Paper 126111. https://doi.org/10.1063/5.0174498 Google Scholar[11] Tang D., Boundary Layer Transition, Science Press, Beijing, 2015, p. 262, Chap. 12. Google Scholar[12] Scott C. J. and Anderson G. E., "Boundary-Layer Transition with Gas Injection," Journal of the Aerospace Sciences, Vol. 25, No. 12, 1958, pp. 791–791. https://doi.org/10.2514/8.7892 LinkGoogle Scholar[13] Marvin J. G. and Akin C. M., "Combined Effects of Mass Addition and Nose Bluntness on Boundary-Layer Transition," AIAA Journal, Vol. 8, No. 5, 1970, pp. 857–863. https://doi.org/10.2514/3.5778 LinkGoogle Scholar[14] Demetriades A., "Hypersonic Viscous Flow over a Slender Cone. III-Laminar Instability and Transition," 7th Fluid and PlasmaDynamics Conference, AIAA Paper 1974-0535, 1974. Google Scholar[15] Jewell J., Leyva I., Parziale N. J. and Shepherd J. E., "Effect of Gas Injection on Transition in Hypervelocity Boundary Layers," 28th International Symposium on Shock Waves,, Vol. 1, Springer, Berlin, 2012, pp. 735–740. Google Scholar[16] Basore K., Selzer M., Wheatley V., Boyce R., Mee D., Capra B., Kuhn M. and Brieschenk S., "Performance Comparison of Distributed Injection Methods for Hypersonic Film-Cooling," Proceedings of the 20th Australasian Fluid Mechanics Conference (AFMC), Australasian Fluid Mechanics Soc., Perth, 2016, pp. 1–4. Google Scholar[17] Miro F. M., Dehairs P., Pinna F., Gkolia M., Masutti D., Reger T. and Chazot O., "Effect of Wall Blowing on Hypersonic Boundary-Layer Transition," AIAA Journal, Vol. 57, No. 4, 2019, pp. 1567–1578. https://doi.org/10.2514/1.J057604 Google Scholar[18] Fedorov A. V., Malmuth N. D., Rasheed A. and Hornung H. G., "Stabilization of Hypersonic Boundary Layers by Porous Coatings," AIAA Journal, Vol. 39, No. 4, 2001, pp. 605–610. https://doi.org/10.2514/2.1382 LinkGoogle Scholar[19] Miró Miró F. and Pinna F., "Effect of Uneven Wall Blowing on Hypersonic Boundary-Layer Stability and Transition," Physics of Fluids, Vol. 30, No. 8, 2018, Paper 084106. https://doi.org/10.1063/1.5043353 Google Scholar[20] Mortensen C. H., "Toward an Understanding of Supersonic Modes in Boundary-Layer Transition for Hypersonic Flow over Blunt Cones," Journal of Fluid Mechanics, Vol. 846, July 2018, pp. 789–814. https://doi.org/10.1017/jfm.2018.246 CrossrefGoogle Scholar[21] Jewell J., Wagnild R., Leyva I., Candler G. and Shepherd J., "Transition Within a Hypervelocity Boundary Layer on a 5-Degree Half-Angle Cone in Freestream air/CO2 Mixtures," 51st AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition, AIAA Paper 2013-0523, 2013. LinkGoogle Scholar[22] Zhao J., Liu S., Zhao L. and Zhang Z., "Numerical Study of Total Temperature Effect on Hypersonic Boundary Layer Transition," Physics of Fluids, Vol. 31, No. 11, 2019, Paper 114105. https://doi.org/10.1063/1.5125116 Google Scholar[23] Wartemann V., Wagner A., Wagnild R., Pinna F., Miró Miró F., Tanno H. and Johnson H., "High-Enthalpy Effects on Hypersonic Boundary-Layer Transition," Journal of Spacecraft and Rockets, Vol. 56, No. 2, 2018, pp. 347–356. https://doi.org/10.2514/1.A34281 Google Scholar[24] Jiang Z. and Yu H., "Theories and Technologies for Duplicating Hypersonic Flight Conditions for Ground Testing," National Science Review, Vol. 4, No. 3, 2017, pp. 290–296. https://doi.org/10.1093/nsr/nwx007 CrossrefGoogle Scholar[25] Liu M. K., Han G. L. and Jiang Z. L., "Experimental Study on the Evolution of Mode Waves in Laminar Boundary Layer on a Large-Scale Flat Plate," Physics of Fluids, Vol. 34, No. 1, 2022, Paper 013612. https://doi.org/10.1063/5.0075710 Google Scholar[26] Liu M. K., Han G. L., Li Z. X. and Jiang Z. L., "Experimental Study on the Effects of the Cone Nose-Tip Bluntness," Physics of Fluids, Vol. 34, No. 10, 2022, Paper 101703. https://doi.org/10.1063/5.0110928 Google Scholar[27] Wang Y., Wang Y. and Jiang Z., "Experimental and Numerical Investigation on the Unsteady Interaction in Longitudinal Stage Separation for Parallel-Staged Two-Body Configuration," Physics of Fluids, Vol. 36, No. 1, 2024, Paper 016116. https://doi.org/10.1063/5.0184293 Google Scholar Previous article Next article FiguresReferencesRelatedDetails What's Popular Articles in Advance Metrics CrossmarkInformationCopyright © 2024 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. All requests for copying and permission to reprint should be submitted to CCC at www.copyright.com; employ the eISSN 1533-385X to initiate your request. See also AIAA Rights and Permissions www.aiaa.org/randp. TopicsAerodynamicsAeronauticsAerospace SciencesAerothermodynamicsBoundary LayersFluid DynamicsHeat FluxThermodynamic PropertiesThermodynamicsThermophysics and Heat TransferVortex DynamicsWind Tunnels KeywordsHeat Flux DistributionAdverse Pressure GradientAerodynamic HeatingHypersonic Boundary LayerWind Tunnel TestsBoundary Layer ControlAcknowledgmentsThis work was supported by the National Natural Science Foundation of China (12132017), the National Key Research and Development Program of China (2022YFB3207000), the Youth Cross Team CAS (JCTD-2022-02), and the Youth Innovation Promotion Association CAS (2020019).Digital Received29 May 2024Accepted31 July 2024Published online3 September 2024

The incident shock attenuation phenomenon in shock tube has received widespread interest because of its inevitable influence on experimental gas properties. However, few studies have investigated the cascading effects of the resulting nonuniformity on shock reflection in shock tunnels before the nozzle. This paper describes a numerical study on the unsteady reflection of a decelerating incident shock driven by a decelerating piston, as a simplification of the complex nonideal factors. The initial decay and uniform parameter distributions are generated based on the non-inertial frame. The results indicate that the existence of the nonuniform area forces the reflected shock wave region to undergo a transition process of attenuation and then stability. The final values of the gas parameters in zone 5 will, therefore, deviate from those given by the traditional relation for an ideal shock tube, which are only dependent on the terminal shock Mach number. This affects the determination of the total temperature T5, which is difficult to measure directly. We discuss the reconstruction of the nonuniform region with the attenuation trajectory of the incident shock wave and find that there is no one-to-one correspondence between this trajectory and the resulting nonuniformity, which introduces additional uncertainties to the predictions. Thus, an analogy method is developed through the multilevel block building algorithm, allowing the total temperature T5 to be determined using the total pressure p5 considering the effect of nonuniformity. Further assessments verify the applicability of this method in cases with multidimensional and viscous interactions.

Air-blowing is one of the techniques for active flow control and thermal protection system of hypersonic vehicles. Introducing air into the hypersonic boundary layer alters the cross-sectional profile of the boundary layer, thereby influencing the boundary-layer transition. This study investigates the active air-blowing control effects on the hypersonic flat-plate boundary layer under various blowing mass flow rates and incoming Mach numbers by solving the Reynolds-averaged Navier–Stokes equations with the Langtry–Menter four-equation transitional shear stress transport model. The study examined alterations in the blowing boundary-layer profiles under two conditions: natural and bypass transition, induced by different blowing flow rates. Blowing significantly alters the sonic line and boundary-layer profile characteristics, triggering blowing oblique shock and causing alterations in the instability mechanisms of the two transition states. A higher Mach number intensifies compressibility effects, stabilizing the boundary layer and leading to an increase in the thickness of the blowing boundary layer and air film.

The interactions between oblique and bow shock waves are significant problems related to the aerodynamic force and thermal performance of hypersonic vehicles, but few studies have considered the dynamic effect of the body's motion on the phenomena. In this work, a numerical study on the oblique and bow shock waves ahead of an elliptic cylinder rotating with a forced-oscillation approach was conducted at Mach 5 by solving the unsteady, two-dimensional Navier–Stokes equations in a non-inertial coordinate system. The hysteresis loops of aerodynamic coefficients were analyzed first, and it was found that the moment is sensitive to rotation. Then, two different hysteresis forms were found at positive and negative angles of rotation (AOR), corresponding to cases with the interference point above or below the wall, respectively. When AOR is positive, the rate-dependent transition hysteresis among various shock interaction types causes the movement of strong flow structures (reflected shock wave, Mach stem, and jet) to always lag behind the body's motion. When AOR is negative, besides the evolution hysteresis of flow structures, two unusual patterns between Edney Types III and VI were observed on different transition paths, which led to very different peak pressures. Also discussed are the driving mechanisms associated with the effect of the subsonic region and the downstream boundary of the interaction zone, as well as the modulating action of the formed virtual Laval flow channel. Additional simulations were performed to study the effect of rotation speed on the transition boundary and the transition structures between Types III and VI.

Hypersonic liquid film cooling technology is to press out the cooling medium through a series of slits or holes, creat a low-temperature cooling film in the boundary layer of the surface of the aircraft to prevent the aerodynamic heating of the aircraft by hypersonic airflow. As an active cooling method, it has great application potential in surface thermal protection of hypersonic vehicle.In this paper, numerical methods and VOF model are used to study the spreading of liquid film at 25km flight altitude and Ma5 airflow. The evolution process and cooling mechanism of liquid film on a flat plate are discussed through the incident velocity, Angle, surface tension and viscosity coefficient of different cooling medium. The results show that under the action of air flow, the liquid film develops downstream to the wall surface, the existence of the liquid film leads to the boundary layer separation, and the continuous liquid film will be broken into liquid blocks at a certain position,and then further broken into droplets. The change of incident conditions and liquid properties will affect the development of the liquid film along the flow direction, which is manifested in the position of the fracture point and the thickness of the continuous liquid film. Within the computational domain set in this paper, the wall heat flow is reduced by 80～95 percent, and the cooling efficiency of the liquid film on the wall varies with the the change of the liquid film morphology.

The research on shock wave/turbulent boundary-layer interactions is mainly limited to calorically perfect gases; little has been reported on the thermochemical non-equilibrium (real gas) effect. This effect is prominent at conditions of high Mach and Reynolds numbers. In this work, a household parallel solver for hypersonic thermochemical non-equilibrium flows with a Reynolds-averaged Navier-Stokes turbulence model is developed in which the coupling of turbulence with vibration and chemistry occurs under a gradient-law assumption. The thermal non-equilibrium is based on Park's two-temperature model, and the chemical non-equilibrium is based on Gupta's 11-species model. The method proposed in this paper is first validated using experimental data, including cases of a laminar cylinder flow at a high-enthalpy condition, a supersonic flat-plate turbulent boundary layer flow, a hypersonic transition flow, and a hypersonic compression corner flow at a low-enthalpy condition. This approach is then applied to assess the hypersonic flow characteristics past the 34 degrees compression corner at a flight height of 30 km. Results show that the joint effects of turbulence and thermochemical non-equilibrium have a significant impact on the flow field organization, wall data, and separation length of the shock wave/boundary-layer interaction. Furthermore, the mechanism of the neck region accompanied by maximum heat flux, wall pressure and skin friction in both laminar and turbulent cases is well-interpreted. This study can be used as a reference tool for the aerodynamic design of future hypersonic vehicles accounting for multi-physics effects.

In this Letter, the dynamic effect of rotation on unsteady shock interaction between an oblique shock wave and bow shock ahead of an elliptic cylinder is numerically investigated. First, a non-deforming grid method based on a non-inertial reference frame is applied for supersonic/hypersonic flow around an arbitrarily moving body. A simulation considering a rotating elliptic cylinder is then conducted at Ma5 with forced-oscillation approach. The results show that the shock interaction structure during rotation has obvious hysteresis compared with the static condition, and transition among different types of shock interaction is also affected by rotation direction and speed. The mechanisms are briefly addressed.

In this research, to study the hypersonic boundary-layer transition, experiments were conducted on a large-scale flat plate with a length of 3.2 m at a zero angle of attack in the hypersonic shock tunnel duplicating flight conditions. Surface-mounted piezoelectric pressure sensors and coaxial thermocouples were, respectively, used to measure the pressure fluctuations and wall heat transfer. The spatial distribution of heat transfer was used to distinguish the transition. Under the test conditions of Ma = 7.0, T0 = 2120 K, and Re∞ = 6.08 × 105 m−1, no transition occurred, and under the test conditions of Ma = 7.0, T0 = 2220 K, and Re∞ = 1.23 × 106 m−1, the transition position was s = 2.06 m. The repeatability of the experiment was found to be good. Furthermore, focus was placed on the spectral and spatial/temporal evolution characteristics of pressure fluctuations in the laminar boundary layer. The experiment captured the three frequency distributions of mode waves in the laminar flow zone. Among the mode waves distributed in the three frequency bands, the low-/high-frequency bands were dominant, and the mid-frequency band exhibited a staged contribution. The amplitude energy percentages of the high- and low-frequency mode waves exhibited opposite trends in both time and space, which means that the disturbance energy will be distributed among the various harmonics in the laminar stage.

In this Letter, hypersonic boundary-layer transition was investigated on a large-scale cone with a height of 3 m and a half-cone angle of 7° at a zero angle of attack in the JF-12 hypersonic flight duplicate shock tunnel. For the same freestream unit Reynolds number, with the increase in the bluntness Reynolds number, the transition Reynolds number has a trend of first increasing and then decreasing, showing a “transition reversal” phenomenon. As the bluntness increased, the high/low-frequency instability waves in the boundary-layer were modulated, which caused the boundary-layer transition to be delayed and then advanced.

采用高精度格式求解二维Navier-Stokes方程,研究了不同飞行高度下超声速来流和射流在后台阶相互作用的流场基本结构.时间推进采用三阶精度Runge-Kutta格式,分别应用五阶精度加权本质无振荡(weighted essentially non-oscillatory,WENO)格式、六阶精度中心差分格式来离散对流项和粘性项,并应用MPI非阻塞式实现并行化.采用两步后台阶模型分别研究了不同高度下超声速后台阶流动、射流的基本结构特征;并进一步组合两种流动,研究了超声速来流/射流组合流动下相互作用的流场结构.通过改变后台阶上方来流条件,模拟了不同飞行高度的环境,研究了其对流场中涡、剪切层、激波等结构的影响.研究结果发现,超声速来流和射流发生相互作用后,在后台阶附近产生回流区,超声速来流的存在会对射流的流场结构产生影响.

In this paper, Navier-Stokes equations were solved with high-order accurate schemes to investigate the basic structure and regularity of the flow field during the interaction of a supersonic jet and a codirectional supersonic incoming flow. A double backward-facing step model was proposed to investigate the interaction between the jet/supersonic incoming flow shear layers. The two shear layers interact to produce a secondary jet. The secondary jet produced by the action has a unique periodicity that is related to the overall oscillation of the shear layer. The secondary jet is generated when the horizontal angle of the jet shear layer reaches a certain value. This paper focused on the analysis and discussion of the periodicity of the secondary jet. When the aspect ratio is different, the period of the secondary jet changes significantly. However, when the static pressure ratio is different, the period of the secondary jet does not change much.