Numerical Investigations of Shock Wave Boundary Layer Interaction in Hypersonic Flows
| dc.contributor.author | John, Bibin | |
| dc.date.accessioned | 2020-09-01T12:29:56Z | |
| dc.date.accessioned | 2023-10-26T09:43:11Z | |
| dc.date.available | 2020-09-01T12:29:56Z | |
| dc.date.available | 2023-10-26T09:43:11Z | |
| dc.date.issued | 2014 | |
| dc.description | Supervisor: Vinayak Kulkarni | |
| dc.description.abstract | Shock wave boundary layer interaction (SWBLI) is a classical example of viscous-inviscid interactions. Occurrence of SWBLI, in hypersonic flowfield, is extremely adverse for the space missions and needs thorough investigations. Enhanced surface heating, alteration in aerodynamic coefficients, vortex shedding etc. are the possible reasons for this consideration. In view of this numerical studies are planned to enhance the understanding about this interaction. However these investigations demand an efficient and accurate solver suitable for high speed flow analysis. Hence an in-house 2D-axisymmetric viscous compressible flow solver "USHAS (Unstructured Solver for Hypersonic Aerothermodynamic Simulations) is developed to achieve the proposed goals. This solver is then validated and verified using the literature reported numerical or experimental test cases. During these studies, it has been noted that the AUSM family of schemes form a better compromise among convergence, stability and accuracy for hypersonic applications. Thus these schemes are preferred onwards for SWBLI studies. The higher order accurate solver, USHAS, is then implemented for ramp induced SWBLI (R-SWBLI) studies. Initial efforts are invested to quantify the effect of various governing parameters such as ramp angle, freestream Mach number, freestream stagnation enthalpy, leading edge bluntness and wall temperature on the interaction. Results obtained from these simulations are compared with qualitative predictions in literature. It is shown that increase in Mach number as well as bluntness of the leading edge reduces the upstream influence resulting in a decrease in the extent of separation. Contrarily, increase in wall temperature and decrease in stagnation enthalpy enhance the separation size. Possibility of turbulent reattachment is also confirmed for leading edge bluntness case through the present simulations upon comparison of the results with the reported experimental data. Numerical studies also illustrate that the total to wall temperature ratio is a better indicator of SWBLI rather than the individual quantities themselves. Subsequently, using the same simulation data, various literature reported correlations to predict incipient separation condition, extent of upstream influence, separation bubble length, plateau pressure and separation pressure, have been critically reviewed and assessed. It has been noted that the correlation for incipient condition should only be used for separated flows modifications are suggested in other correlations to widen their range of applicability. The widely accepted control technique, provision of leading edge bluntness, is investigated to assess its capability to delay or avoid the separation. Various dependent parameters like sonic height, boundary layer edge Mach number, entropy layer, boundary layer thicknesses etc. | en_US |
| dc.identifier.other | ROLL NO.09610311 | |
| dc.identifier.uri | https://gyan.iitg.ac.in/handle/123456789/1658 | |
| dc.language.iso | en | en_US |
| dc.relation.ispartofseries | TH-1295; | |
| dc.subject | MECHANICAL ENGINEERING | en_US |
| dc.subject | MECHANICAL ENGINEERING | |
| dc.title | Numerical Investigations of Shock Wave Boundary Layer Interaction in Hypersonic Flows | |
| dc.type | Thesis |