Computation of Hypersonic Flows with Lateral Jets Using k-ω Turbulence Model (original) (raw)
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Hypersonic Vehicles - Applications, Recent Advances, and Perspectives [Working Title], 2022
The goal of this work is to give optimum aerothermal solutions for thermal protection of the nose wall of space shuttles during atmospheric reentry, where the air flow is hypersonic, nonequilibrium reactive flow (vibrational and chemical) behind detached shock waves, it’s governed by Navier–Stokes equations with chemical reaction source terms, and modelled using five species (N2, O2, NO, N, O) and Zeldovich chemical scheme with five reactions. This study which simulates the flow using the software Fluent v.19 focuses on the comparison between three protection techniques based on the repulsion of the shock wave, the first is geometric, it consists in introducing a spike that makes the right shock move away from the nose of the shuttle, this allows the endothermic physicochemical processes of dissociation and ionization to absorb heat, the second technique is based on an opposite jet configuration in the frontal region of the nose, this jet allows to push the strong shock, and consequ...
Numerical investigation of drag and heat reduction in hypersonic spiked blunt bodies
Heat and Mass Transfer, 2013
In this investigation, the effects of spike as retractable drag and aerodynamic heating reduction into the reentry Earth's atmosphere for hemispherical body flying at hypersonic flow have been numerically studied. This numerical solution has been carried out for different length, shapes and nose configuration of spike. Additional modifications to the tip of the spike are investigated in order to obtain different bow shocks, including no spike, conical, flat and hemispherical aerodisk mounted. Unsteady compressible 3-D Navier-Stokes equations are solved with k-x (SST) turbulence model for a flow over a forward facing spike attached to a heat shield for a free stream Mach number of 6. The obtained numerical results are compared with the experimental ones, and the results shows acceptable verification. This analysis shows that the aerodisk is more effective than aerospike. The designs produced 60 and 15 % reduction in drag and wall temperature responses, respectively. List of symbols C f Surface skin friction coefficient C P Specific heat at constant pressure (J/kg K) C p Static pressure coefficient D Payload shroud diameter (m) e Specific energy (J/kg) F, G Inviscid flux vector H Source vector M Mach number p Static pressure (N/m 2) Pr Prandtl number q Heat flux (W/m 2) Re Reynolds number R, S Viscous flux vector t Time (s) U Mean stream wise velocity (m/s) u, v Velocity components (m/s) W Conservative variables in vector form x, r Coordinate directions (m) Greek symbols l Molecular viscosity (kg/m s) q Density (kg/m 3) r rr , r xx Normal stress tensor (N/m) s xr Shear stress tensor (N/m) Subscripts w Wall 1 Free stream condition 1 Introduction Researches on effects of forward faced spike on hypersonic flight have been started since the late 1940s [1]. The primarily concern has been accorded to drag characteristics due to spike length, spike nose geometry, forward body geometry and relative spike diameter have been studied. The drag exerted on a body in hypersonic flow is an important issue of high speed aerodynamics. For reentry vehicles, it is important to control the deceleration of a reentry capsule by
Investigation on thermal protection and drag reduction by lateral jet in supersonic flows
Proceedings of the6th International Conference on Mechatronics, Materials, Biotechnology and Environment (ICMMBE 2016), 2016
A numerical code was developed to investigate the characteristics of drag and heat reduction by the lateral jet in supersonic flows, the Reynolds-average Navier-Stokes (RANS) equations was solved by using high resolution upwind scheme AUSMPW+, three order MUSCL reconstruction method and k-ωSST turbulence model with solving the unsteady heat transfer equations. The code was validated by an experimental case. The computed results indicate that: the numerical code can effectively capture the flow features, and the pressure and Stanton number on blunt body surface are significantly decreased with the addition of the lateral jet. The lateral jet was proved to be an effective method to reduce drag and heat.
Aerothermodynamics of Generic Reentry Vehicle with a Series of Aero spikes at Nose
Re-entry of a blunt nosed body is one of the most intriguing problems in any space program. Especially in the light of various space tourism possibilities, there are many works concerning re-entry of commercial blunt nosed space vehicles. In this paper, a generic blunt body re-entry model represented by a hemisphere- cylinder, fitted axisymmetrically with an aerodisk aerospike at the nose is investigated numerically with commercially available control volume based a density based, axisymmetric flow solver ANSYS Fluent 14.0. The scaled down re-entry model has a base diameter of 40 mm and an overall length of 100 mm. A 6 mm diameter aerospike fitted axisymmetrically at the nose has a hemispherical cap from which another aerospike of 4 mm diameter protrudes which again has a hemispherical cap. A two dimensional compressible, axisymmetric Navier Stokes Equations are solved for a hypersonic flow of thermally perfect air with free stream conditions of Mach No. 21.08 and a static pressure, altitude and temperature of 37.362 Pa, 55.842 km and 258.02 K respectively. The results are compared with that of re-entry capsule without aerospike. Among the cases investigated, the spiked blunt body having two aerospikes in series with lengths l1 and l2 equal to 30 and 20 respectively shows a 38 percent reduction in peak reattachment heat flux and 25 percent reduction in drag and thus stands as a prospective case for blunt body nose configuration for hypersonic flight.
Computational Analysis of Hypersonic Flow Field Over Re-Entry Bodies
The main objective of this paper is to conduct computational analysis of a typical re-entry vehicles. Computational fluid dynamics is used to obtain the flow field that develops around reentry capsulesflare shape in hypersonic flow. This problem is of particular interest since it features most of the aspects of the hypersonic flow around a planetary entry vehicle. The region between the cone and the flare is particularly critical with respect to the evaluation of the surface parameters. Flow separation induced by the shock wave / boundary layer interaction, with subsequent flow reattachment, that can dramatically enhance the surface heat transfer. The exact determination of the extension of the recirculation zone is particularly a difficult task for numerical simulation. The analysis is carried out for turbulent flow and standard flow properties available for Re-entry capsules in the literature using full Navier-Stokes solver for different Mach numbers. The study is carried out using the commercial CFD code. The computational analysis of Re-entry bodies involving relative enthalpy, wall fluxes, wall shear stress, and pressure coefficient at various locations of the capsules are presented.
Numerical Study of Flow Augmented Thermal Management for Entry and Re-entry Environments
25th AIAA Applied Aerodynamics Conference, 2007
Tremendous amounts of heat and drag loads occur during entry and re-entry into planetary and earth atmospheres have posed grave challenges on maintaining structure integrity of space exploration vehicles. Though various thermal protection systems (TPS) have been employed to manage the heat loads, both localized and transient spikes at stagnation points, the use of TPS can substantially increase the weight of the vehicle. Hence, various concepts, such as aeroassisted orbital transfers and aerobraking, have been designed to mitigate the high heating rates such that the TPS weight can be minimized. Among those concepts, the use of a flow augmented thermal management system for entry and re-entry environments has shown some promises in reducing heat and drag loads. This concept relies on jet penetration from supersonic and hypersonic counterflow jets, that could significantly weaken and disperse the shock-wave system of the spacecraft flowfield to reduce wave drag and aerothermal loads. Greatly reducing wave drag and aerothermal loads significantly enhances aerothermal performance, allowing thinner or much lighter TPS to be used, which translates into spacecraft weight and cost savings. Other benefits include better aerodynamic efficiency and improved down-range and cross-range maneuverability. There are two jet penetration modes involved in a supersonic/hypersonic flow interacting with counterflow jets: short penetration mode (SPM) and long penetration mode (LPM) interactions. Previous studies have shown that the LPM jet significantly increases the shock stand-off distance, thus reducing the strength of bow shock, which leads to a reduction in wave drag. The LPM jet acts as, in essence, a "pencil" of fluid with high dynamic pressure, penetrating into the incoming freestream, to attenuate the shock system. Though the function of the LPM jet has been demonstrated in the previous study, further experimental and computational analyses through trade studies are required to determine the optimum operating conditions of the LPM jet.
Investigation and recent developments in aerodynamic heating and drag reduction for hypersonic flows
Heat and Mass Transfer, 2018
Investigation on new methods of drag reduction and thermal protection for hypersonic velocities are proposed based on the new technologies for shock reconstruction. In principle, a blunt vehicle flying at high speeds generates a strong bow shock wave ahead of its nose, which is responsible for the high drag and aero heating levels. There have been a number of efforts devoted towards reducing both the drag and the aero heating by modifying the flow field ahead of the vehicle's nose. An introduction to the philosophy and recent development in hypersonic aerodynamic heating and drag reduction techniques are presented. These techniques are classified in four major group. Geometrical, mass injection, energy deposition, and magneto aerodynamic techniques. In this review, these new techniques and investigation of the philosophy and development procedure of these techniques are brought to the table and then the effects of each method on drag wave and aerodynamic heating reduction is shown. Geometrical techniques such as structural spike and aerodisk, cavity, multi-row disk (MRD) use to drag reduction. Mass injection techniques like arrays of micro jets, spike and jet, heat addition and plasma injection are very useful to aerodynamic heating reduction. Energy deposition techniques by using laser and plasma can reduce aerodynamic heating and wave drag of hypersonic flows. Magneto aerodynamic technique can reduce drag and aerodynamic heating, by the addition of the magneto technology in hypersonic flows. The present paper is devoted to surveying these studies and illustrating the contributions of the authors in this field. Not only do the paper criticize the previous investigations but also raises some of the areas in the field that need further investigations.
System Studies on Active Thermal Protection of a Hypersonic Suborbital Passenger Transport Vehicle
Aerodynamic heating is a critical design aspect for the development of reusable hypersonic transport and reentry vehicles. The reliability in terms of thermal resistance is one of the major driving factors with respect to the design margins, the mass balance and finally the total costs of a configuration. Potential designs of active cooling systems for critical regions such as the vehicle nose and leading edges are presented as well as preliminary approaches for their impact on the total mass. The visionary suborbital passenger transport concept SpaceLiner is taken as a reference vehicle for these studies. Covering the whole flight regime from subsonic to Mach numbers of more than 20, this vehicle creates high demands on the thermal protection system. Part of the work was performed within the DLR research project THERMAS.
Computational study of flow over Fire II Re-entry Vehicle at various Supersonic Speeds
International Journal of Advance Research and Innovative Ideas in Education, 2021
Re-entry modules are blunt bodies that are designed in such a way that they can withstand the extreme temperature loads during atmospheric re-entry. Designing a re-entry vehicle requires dependable data on the aerodynamic flow properties. After hypersonic re-entry, the vehicle has to decelerate at supersonic speeds until the parachutes are deployed. Examining the flow properties over the vehicle at supersonic speed is important of which temperature serves as critical data for a successful recovery operation. Hence this paper is presented to understand the variation in temperature, as the re-entry vehicle decelerates from Mach no. 3 to 1.2. Temperature variation along with pressure distribution is obtained computationally for freestream Mach numbers 1.2, 2 and 3 with appropriate boundary conditions using Ansys FLUENT. The results also describe the flow field around a capsule in supersonic flow and understand the importance of shock waves which influences the aerodynamic forces on the capsule. The analysis showed the flow-field features such as bow shock wave on the frontal heat shield surface and expansion fan on the shoulder terminating region around the re-entry capsule. This research will assist spacecraft design engineers to analyse the flowfield and choose the appropriate materials. Further, the analysis performed can aid in the designing of parachutes that are used to decelerate the re-entry vehicle for a safe landing.
Numerical analysis of hypersonic flows around blunt-nosed models and a space vehicle
Aerospace Science and Technology
This work addresses the problem of the aerothermodynamics of hypersonic nonequilibrium flows over blunt nosed models and space vehicles with rarefaction effects. First, the in-house Navier–Stokes solver, UNIC-UNS code, with the slip boundary condition and finite-rate chemistry is used to simulate the hypersonic flows over a blunt nosed model and the simplified European eXPErimental Re-entry Test-bed (EXPERT) model V4.4. Next, hypersonic flows over the whole EXPERT 3D model, which correspond to the expected descent trajectory with allowance for rarefaction and thermochemical nonequilibrium are simulated. By comparing with the Direct Simulation Monte Carlo (DSMC) method, it is observed that the UNIC-UNS code is reliable in simulating hypersonic flows with rarefaction and thermochemical non-equilibrium effects. A detailed analysis of the aerothermodynamics for EXPERT for a wide range of flow regimes is also provided by utilizing the numerical flow visualization. The present numerical s...