1. Introduction

During the design phase of launchers, the aerodynamic characterization represents a fundamental contribution. Usually, it is accomplished by means a hybrid approach encompassing wind tunnel testing (WTT) and computational fluid dynamics (CFD) investigations [1]. This combined design approach (i.e., WTT and CFD analyses) is extremely reliable in providing high quality data as input for launchers' sizing, performance evaluations, control, and staging dynamics [2]. Indeed, launcher aerodynamics focuses on the assessment of the pressure and skin friction loads the atmosphere determines over the vehicle surface [3]. As well known, these loads result in a global aerodynamic force that acts at the aeroshape center of pressure (CoP) which generally does not coincide with the vehicle center of gravity (CoG) [4]. As a result, the related aerodynamic moment acting at the CoG can lead to a stable or unstable behavior of the launcher to account for in the control software [5]. Moreover, the analysis of the flowfield past the launcher is also fundamental to address the effects of aeroshape's structures

© 2018 The Author(s). Licensee InTech. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/by/3.0), which permits unrestricted use, distribution, and reproduction in any medium, provided the original work is properly cited.

© The Author(s). Licensee InTech. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/by/3.0), which permits unrestricted use, distribution, and eproduction in any medium, provided the original work is properly cited.

and protrusions. Indeed, aeroshell steps and gaps determine local pressure (and convective heat flux) overshoots all along the ascent trajectory [6]. This assessment is fundamental for launcher sizing and thermal protection design activities [7].

can impinge on the launcher aeroshape and cause local pressure and heat flux overshoots, well

Launcher Aerodynamics: A Suitable Investigation Approach at Phase-A Design Level

http://dx.doi.org/10.5772/intechopen.70757

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On the other hand, SWBLI occurs, for instance, when the shock resulted from the SSI meets the

As a result, SSI and SWBLI demand accurate prediction for a reliable and affordable aero-

In this framework, the results of the computational analysis of the flowfield past a double wedge test bed are reported and discussed in detail. This configuration, in fact, is a benchmark as it presents unique flow patterns typical of SSI and SWIBLI. In particular, the experiment of Swantek and Austin was selected and numerically rebuilt [9]. The test bed geometry is shown in Figure 2. It is a double wedge with θ<sup>1</sup> = 30 and θ<sup>2</sup> = 55 where the lengths of the first and

Along with the center of the model 19 coaxial thermocouple gauges at 16 different streamwise locations are mounted. Therefore, several experimental data exist for numerical-to-experimental comparisons. The test campaign was performed by using high enthalpy air at the free-

The numerical rebuilding was carried out by means of a steady-state two-dimensional Reynolds-averaged Navier-Stokes (RANS) simulation performed with the commercial CFD tool Fluent. Air was modeled with a five species chemistry mixture (N2, N, O2, O, NO) in

Parameter M7\_8 Stagnation enthalpy (MJ/kg) 8.0 Mach 7.14 Static temperature (K) 710 Static pressure (kPa) 0.780 Velocity (m/s) 3812

) 0.0038

/m) 0.435

launcher wall, thus promoting boundary layer separation and transition.

second face are L1 = 50.8 mm and L2 = 25.4 mm, respectively.

in excess of those occurring at stagnation points.

thermal design of launcher vehicles.

stream conditions summarized in Table 1.

Figure 2. Test bed configuration with quotes.

Table 1. Free-stream conditions of experiment.

Density (kg/m<sup>3</sup>

Unit Reynolds number (106

With this in mind, the present research effort describes typical aerodynamic analyses performed at Phase-A design level [8]. Indeed, engineering-based analyses are carried out by exploiting local surface inclinations methods. After that, fully three-dimensional steady-state CFD analyses have been addressed to feed launcher aerodynamic design in the range between Mach 0.5 and 5.

Nevertheless, this chapter opens focusing attention on the assessment of the reliability of the present numerical design approach. Indeed, a CFD validation study was undertaken in order to highlight the capability of this CFD approach in assessing some critical aerothermal design issues, namely shock-shock interaction (SSI) and shock wave boundary layer interaction (SWIBLI), of vehicle aeroshapes flying at hypersonic speed, like launchers.

Finally, note that numerical flowfield analyses are performed with FLUENT code and perfect gas flow model.
