Computational flowfield analyses of hypersonic problems with reacting boundary layer
Introduction
In the field of aerospace research, numerical simulations play an important role as they allow to significantly reduce the number of in flight and plasma wind tunnel (PWT) experimental test campaigns. Numerical simulations for flight vehicle design lead to great benefits; for example, vehicle with enhanced aerodynamics or provision of much more accurate vehicle aerodynamic and aerothermodynamic databases for reliable flight missions. The compressible Navier–Stokes equations represent the most sophisticated mathematical model of Newtonian fluid flows and the numerical techniques to solve these equations are very complex and time consuming.
Generally speaking, a flow is defined hypersonic when the Mach number (Ma) is larger than 5. This is a conventional choice, as at this value there is not a significant change in the mathematical form of the governing equations, in contrast to what happens at where, ranging from subsonic regime () to supersonic one (), the equations change from elliptic to hyperbolic form, respectively. In the hypersonic regime (), we observe only an experimental evidence of physical phenomena that happen at Ma larger than 5: in particular, the very high temperature reached in the shock layer, or the interaction between shock wave and boundary layer around the vehicle. At these high temperatures ( K), the air cannot be considered as a perfect gas, but rather as a mixture of reacting species, since diatomic molecules of oxygen and nitrogen start to dissociate. Moreover, species vibrational excitation must be considered, because the molecules are characterized not only by translational and rotational, but also by vibrational degrees of freedom. As the ratio between the specific heats () depends on the number of species active degrees of freedom, it is evident that when the temperature increases, the value of cannot anymore be considered as a constant (perfect gas hypothesis). The main consequence of these modifications is that the Navier–Stokes equations for a perfect gas result inadequate for modeling an high temperature flow, since a much more sophisticated mathematical model is needed.
In this paper, we provide a mathematical model able to describe hypersonic high temperature flow fields and the numerical methodology used. Then, some practical applications of the proposed methodology are discussed. The first one is the simulation of the flow inside the conical nozzle of SCIROCCO facility, located at CIRA. The second application deals with a blunt sphere cone configuration, namely ELECTRE, that is a standard reference model to study non-equilibrium hypersonic flow past blunt-body. The third application refers to a capsule configuration for crew reentry vehicle (CRV) to support system of the international space station (ISS). As last application, the numerical activities performed to evaluate the flowfield around a Wing-body demonstrator suitable for in flight experimental researches were reported.
Section snippets
Mathematical formulation and numerical solution
The mathematical formulation describing a flowfield about an hypervelocity vehicle deals with balance equations for a multi-species chemically reacting gas mixture. In fact the strength of the shock wave produced ahead of a vehicle travelling at hypersonic speed () suddenly increases the temperature of air surrounding the vehicle. Hence at high temperature ( K) the air cannot be anymore considered as a perfect gas, but rather as a mixture of several reacting and vibrating species, and
Scirocco PWT nozzle
The first practical application of the described methodology, is the simulation of the flow inside the nozzle of SCIROCCO PWT facility, shown in Fig. 1. SCIROCCO is a plasma hypersonic wind tunnel located at CIRA. It is based on an electric arc heater, with a maximum power of 70 MW. By means of this facility it is possible to simulate experimentally the aerothermal environment that a space vehicle experiences along an entry into Earth’s atmosphere. Each simulation is identified by the couple of
ELECTRE test article
ELECTRE test article consists of a blunt conical surface with total length of 0.4 m, semiaperture cone angle of 4.6°, and hemispherical nose with radius of 0.035 m. It was tested in flight and in wind tunnel, becoming a standard reference model to study non-equilibrium hypersonic flow past blunt-body configurations [13]. The computational grid employed (Fig. 4) consists of cells with a minimum normal wall spacing of m. Here we analyze the results obtained simulating the flow in one of
LEO manned reentry of an Apollo-shaped capsule
As further application, we have considered a manned reentry from low Earth orbit (LEO) of a crew reentry vehicle (CRV). An Apollo-shaped vehicle configuration is assumed for the CRV [19], [20], which measures about 5 m in diameter, with a nose radius of 6.05 m, a sidewall angle of 33°, and an overall height of 3.8 m (Fig. 6). Entry analysis from LEO orbit was provided for a capsule vehicle gross weight of about 9 ton. For this value of vehicle mass, two descent trajectories were calculated: a
Wing-body configuration
As last application, the numerical activities performed to evaluate the flowfield around a Wing-body demonstrator suitable for in flight experimental researches were reported. The flight demonstrator features a conventional Wing-body configuration with a sharp-leading-edges double-delta planform as basic shape (Fig. 13). CFD analyses of the vehicle along the descent flight path refer to freestream conditions summarized in Table 2. The simulation is performed at laminar flow conditions (due to
Conclusions
An overview of numerical technique able to deal with high temperature gas effects in hypersonic flowfield has been discussed. Numerical investigations of classical hypersonic problems and comparison with experimental data have been provided. Four applications, typical of hypersonics, were considered. Real gas effects on the design of a crew reentry vehicle were taken into account. A comparison is made between CFD computations for perfect gas and reacting gas mixture, in order to highlight real
References (22)
Approximate Riemann solvers, parameter vectors and difference schemes
J. Comp. Physics
(1981)Effects of surface catalytic activity on stagnation heat transfer rates
AIAA Journal
(1973)Hypersonic and High Temperature Gas Dynamics
(1989)An Introduction to Numerical Analysis
(1989)- et al.
Real-gas flowfields about three-dimensional configurations
Journal of Spacecraft and Rockets
(1985) - J.J. Bertin, The effect of protuberances, cavities, and angle of attack on the wind-tunnel pressures and heat-transfer...
Hypersonic Aerothermodynamics
(1994)- R.N. Gupta, J.M. Yos, R.A. Thompson, K.P. Lee, A review of reaction rates, and thermodynamic and transport properties...
- et al.
Thermo-chemical nonequilibrium effects on the aerothermodynamics of aerobraking vehicles
Journal of Spacecraft and Rockets
(1993) - et al.
it Molecular Theory of Gases and Liquids
(1963)
Cited by (6)
AERODYNAMIC AND AEROTHERMODYNAMIC ASSESSMENT OF A LIFTING-BODY RE-ENTRY VEHICLE
2022, 33rd Congress of the International Council of the Aeronautical Sciences, ICAS 2022Multidisciplinary optimization andcfd validation for the design of anext generation re-entry glider
2021, Advances in Engineering Research. Volume 43OPTIMAL DESIGN OF A NEXT GENERATION HIGH-LIFT REUSABLE RE-ENTRY VEHICLE
2021, 32nd Congress of the International Council of the Aeronautical Sciences, ICAS 2021Elaboration of collisional-radiative models for flows related to planetary entries into the Earth and Mars atmospheres
2013, Plasma Sources Science and TechnologyNumerical simulation and experimental validation of a hypersonic flow for numerical modulation of re-entry phenomena prediction using adaptive mesh refinement
2013, International Journal of Computational Methods and Experimental MeasurementsGlobal rate coefficients for ionization and recombination of carbon, nitrogen, oxygen, and argon
2012, Physics of Plasmas