**3. Plasma preheating methodology**

The model in the present work can be likened to the Euro-model of the European Space Agency (ESA) which is a flat-faced cylinder of 50 mm in diameter with rounded corners of 1 mm radius, positioned at a 0° angle of attack to the incoming flow. Preheating of the graphite disc was achieved with plasma [30, 31] generated by a DC current [32, 33] between a tungsten electrode and the back (downstream) side of the disc. **Figure 2** is a sectional view of the model which illustrates the heat transfer processes from the hot plasma to the disc. The orientation of the tungsten inert gas (TIG) is centralised in the model to enable even thermal spread from the centre to the edges. This made it possible for the probe model to assume an axisymmetric orientation.

The disc was set up and heated for 15 seconds. The flow was initiated after 15 seconds of heating, manually timed. The heating remained on during the flow and for about 0.5 seconds after the flow stopped. The present work adopts heatshields of a constant thickness of 2 mm and the thickness was considered suitable for ablation experiments [27]. Using graphite material and a current rating of 400 A, it took about 15 seconds for the plain disk-sample to attain a steady state. Using this current and invoking graphite material properties into FEA simulation with Ansys [29], the results from simulations were found to have a good agreement with experiments as shown in **Figure 3**. The first 15 seconds was for heating, followed by the flow which lasted for about 0.5 seconds, and then finally cooling due to power cut-off from supply. The Schlieren technique based on the principle of changing densities in the gas was used to identify the establishment of the bow shock and therefore the commencement of flow [35]. Obtaining the plenum pressure from pressure survey array, run-times were matched with Schlieren images to identify when flow starts [27]. The high-speed camera was set at a frame-rate of 2500 fps and frames at the specific points during the flow are shown in **Figure 3**.

**Figure 4** illustrates the flow-dynamics and the associated aerothermodynamic gradient along the surface, where the source of heat flux for the heatshield is the plasma inside the probe. The temperature is driven by the plasma at the backside of experimental sample, while the aerodynamic flow is driven by the forced convection from hypersonic impulse facility at the frontside of sample (heatshield specimen). The aerodynamic flow-velocity at the surface of the experimental

**Figure 3.** *Transient behaviour of stagnation surface temperatures with heatshields of 2 mm thickness [28].*

#### **Figure 4.**

*Schematic illustration of surface aerothermodynamic flow properties [36].*

sample increases from the stagnation point to the edges, while the surface temperature decreases from the stagnation point to the edges. While the aerothermodynamic flow (cooling of sample) occurs at the front, the plasmadynamic flow (heating of sample) occurs at the backside of the disk. The plasma zone describes the region occupied by the plasma [36]. At steady state conditions, the inert gas flows through the shroud, gain some heat energy from the centralised hot tungsten electrode, then experiences a drastic rise in enthalpy as it passes through the hot plasma towards the heatshield, before finally exiting via the vent as shown in **Figure 5**.

*Plasma Preheating Technology for Ablation Studies of Hypersonic Reentry Vehicles DOI: http://dx.doi.org/10.5772/intechopen.100129*

**Figure 5.** *Test conducted at Mach 4.5 to validate plasma preheating technology [28].*

The aerothermodynamic flow properties for the plain disk of radius 25 mm were simulated using Ansys Fluent CFD. The CFD results using an axisymmetric graphite sample, showed a good temperature distribution for ablation rate experiments in hypersonic impulse facilities. The temperature is driven by the plasma at the backside of experimental sample, while the flow-field is driven by the hypersonic impulse facility at the frontside of sample [36]. **Figure 6** shows the density variations from Mach 0 to 6, using 2D axisymmetric simulations. The density gradually increased from no-flow conditions to Mach 6 hypersonic flow. Numerical and experimental analysis at Mach 4.5 have been presented extensively by the author in other publications [28, 29].

The reaction species were simulated by using 21% O2 and 79% N2 at the hypersonic inlet, to flow over heated graphite surface [29]. The pressure gradient in **Figure 7a** shows the bow shock standoff distance, while **Figure 7b** shows the contour result of Carbon II oxide species using 2D axisymmetric simulation in Mach 4.5 flow. The ablation rate gradually decreased from stagnation point to the edges of the disk. Along the stagnation line of **Figure 7a**, the upstream edge of the experimental bow shock was about 15 mm from the surface, very similar to the bow shock position described in **Figure 5**. The numerical and experimental results from Mach 4.5 were very similar to that of Mach 6.

Flow properties associated with the present work were obtained from CFD simulations using ANSYS Fluent. **Figure 8** shows the simulation results using a density-based solver for the physical parameters along the stagnation line for a plain disk sample [29]. Surface temperatures and boundary conditions in the CFD were set using measurement information from experiments. **Figure 8** indicates that the temperature has not changed significantly until within 1 mm from the wall, so the concentration of products was not entirely driven by temperature alone. The Mach 4.5 flow parameters were compared with that of Mach 6. The values along the stagnation line were roughly the same. The origin of the horizontal-axes in these figures is at the surface of the disk. In the case of Mach 4.5, the pressure increased from 400 Pa to 18.6 kPa across the shock in the flow direction. The gas density increases from 0.045 kgm<sup>3</sup> to 0.22 kgm<sup>3</sup> across the shock, and continues to rise to a
