Electron Transpiration Cooling for Hypersonic Vehicles


Overview

Hypersonic vehicles require a sharp leading edge in order to increase the lift-to-drag ratio. However, the aerodynamic performance gains offered by sharp leading edges come at the cost of intense localized heating rates. Current thermal protection system materials may not be appropriate because they cannot sustain the large heating rates or are brittle. An approach that has been recently proposed involves using thermo-electric materials at the leading edges of hypersonic vehicles [1]. When exposed to high convective heating rates, these materials emit a current of electrons that leads to a transpiration cooling effect at the surface of the vehicle. This phenomenon is known as thermionic emission and occurs when the thermal energy given to the electrons is greater than the binding potential of the surface material. Electron transpiration cooling (ETC) complements radiative cooling from the surface to balance with the convective heating from the flow. This work focuses implementing ETC into the CFD code LeMANS to investigate the challenges and benefits it can provide to hypersonic flight including the effect of ETC on:

  1. surface temperature
  2. surface heating
  3. plasma physics

Recent Results

A test case was created with geometry representative of the leading edge of a hypersonic vehicle as shown in Fig. 1. The freestream conditions considered are freestream velocities of 4, 6, and 8 km/s (Mach 13 – 26) at an altitude of 60 km. The material work function is 2.0 eV and emissivity is assumed to equal 1.0.

Figure 1: Test case geometry

The flowfield features for the conditions investigated in this test case are highlighted in Fig. 2, which shows temperature contours. The top half corresponds to the results obtained without ETC, and bottom half is for the results obtained with ETC. The figure shows that ETC has an overall minimal effect on the flowfield features.

Figure 2: Temperature contours: without ETC (top), with ETC (bottom)

The temperature and heat transfer distributions along the vehicle surface are presented in Figs. 3. Note that the distance along the leading edge is normalized by the nose radius. Figure 3 illustrates the surface cooling effect due to ETC. For instance, the temperature at the stagnation point decreases from approximately 2600 K for the 6 km/s case without ETC to 1600 K with ETC, which corresponds to a 40% reduction. The drop in surface temperature, however, causes an increase in convective heat transfer to the vehicle as can be seen in Fig. 3b. The reason for this is that the decrease in the surface temperature increases the temperature gradient at the wall, which according to Fourier’s law increases the convective heat transfer to the vehicle. The increase in the heat transfer is localized to a small region near the stagnation point.

(a) Surface temperature

(b) Convective heat transfer

Figure 3: Surface temperature and heat transfer profiles

Figure 4 shows the effect ETC has on radiative cooling. Without ETC, the radiative cooling balances with the convective heat transfer (Fig. 3b). But for the case with ETC, radiative cooling is greatly decreased and ETC becomes the basis of cooling, especially for the higher freestream velocities.

Figure 4: ETC vs. radiative cooling

Investigators

Kyle Hanquist

Acknowledgments

This material is based upon work supported by the Lockheed Martin Corporation.

Recent Publications
References
  1. Uribarri, L. A. and Allen, E. H., "Electron Transpiration Cooling for Hot Aerospace Surfaces," AIAA Paper 2015-3674, July 2015.