**8. Mars entry and descent flight**

#### **8.1 Entry with no bank modulation**

Mars lifting entry is studied assuming the initial conditions expressed by Eq. (3):

$$W\_{\epsilon} = \text{3484 } m/s; \chi\_{\epsilon} = -0.\text{\textdegree{ $ } \text{$  }}^{\text{\textdegree}}; \mu\_{a} = \text{0\textdegree{ $ }}^{\text{\textdegree}}; a\_{\epsilon} = \text{20\textdegree{$  }}^{\text{\textdegree}}; \text{30}^{\text{\textdegree}}; \text{45}^{\text{\textdegree}} \tag{11}$$

## *Hypersonic Vehicles - Applications, Recent Advances, and Perspectives*


**Table 2.** *Loads constraints.*

In **Table 2** the structural and aerothermal constraints which define the entry corridor are shown.

The descent analysis is first performed at zero bank-angle to compute the nominal vehicle trajectory. With *<sup>μ</sup><sup>a</sup>* <sup>¼</sup> 0° the downrange distance depends only on aerodynamic efficiency (i.e., AoA modulation). In **Figure 5a** the effect of AoA on the entry trajectory is shown.

As α increases, the drag coefficient rises and the descent trajectory sinks more, thus moving to lower altitudes, see **Figure 5b**.

Looking at **Figure 5b** we see that with α = 20° the vehicle reaches M<sup>∞</sup> = 2 at an altitude of about 10 km higher than the case at α = 45°. At α = 20°, the spacecraft features a higher hypersonic aerodynamic efficiency. Therefore, higher *L=D* values appear more convenient for Mars missions to limit the terminal landing speed allowing at the same time a greater total range.

**Figure 5c**-**5d** confirms the advantage of flying with a high L/D. Long entry time (one-hour order-of-magnitude) could also help with the landing spot customization. However, for lifting bodies, increasing *L=D* leads to an increase of flight time (see **Figure 5d**), and integrated heat loads, see Eq. (5).

Finally, **Figure 5e**-**5f** show the effect of *α* on the peak heat flux rate and on its time history. At *<sup>α</sup>* <sup>¼</sup> <sup>45</sup>° the peak heat flux is 90 kW/m<sup>2</sup> , while decreases to 50 kW/ <sup>m</sup><sup>2</sup> at *<sup>α</sup>* <sup>¼</sup> <sup>20</sup>° . Therefore, the thermal peak is reduced, and non-ablating (re-usable) materials can be adopted. On the other hand, the longer flight time requires high emissivity materials to decrease heat transfer by conduction. In **Table 3** it is shown the total energy absorbed during entry.

In **Figure 5c** it is shown that for *<sup>γ</sup>* ¼ �0*:*3° entry times are of the order of 10<sup>4</sup> s. To reduce entry flight time, steeper flight-path angles can be considered, see **Figure 6**.

Assuming *<sup>γ</sup>* ¼ �2*:*4°, entry times are of the order of 10<sup>3</sup> s which can be compared to Earth re-entry flight time. However, progressively decreasing *γ* the peak heat flux rises (see values attained at *γ* ¼ �2*:*4°Þ. Therefore, a trade-off study between entry time and flight path-angle suggests the appropriate value related also to thermal insulation material capabilities.
