Total System Thermal Integration

When discussing propulsion, hypersonic flight or atmospheric entry the question of cooling must be examined in the context of the total energy management or integration. In the case of the SR-71 the aerodynamic heating was mostly absorbed by the structure, and the surface ran at radiative equilibrium temperature. So the SR-71 was a hot structure vehicle and therefore it required a material that maintained its strength at high temperature (i.e., in the 660°C range) and that was beta-titanium. The thermal energy had to be removed from the crew compartment and equipment bays. That thermal energy plus the thermal energy rejected by the engine was transferred to the fuel. Discussions of the SR-71 design state that the fuel temperature entering the engine was over 600° C. In this case all of the thermal energy was discarded as hot fuel and that hot fuel provided no useful work or engine thrust. With a high-temperature hydrocarbon as fuel this was a rational approach as there was hardly any option to extract the recovered energy from the liquid hydrocarbon.

With a fuel that is a good heat transfer medium, the structural concept is unlike the SR-71 hot structure, but more like a cold structure protected by a combination of

Thermal recovery panels

Thermal recovery panels

Figure 4.17. System thermal integration.

Figure 4.18. Closed cycle heat pump (after Ahern) and combustor fuel injection.


Figure 4.18. Closed cycle heat pump (after Ahern) and combustor fuel injection.

radiation shingles, radiating about 95% of the aerodynamic heating back to space, and a structural thermal management system that converts about half of the thermal energy entering the airframe into useful work and thrust. Figure 4.17 illustrates a systems level thermal integration approach [Ahern, 1992]. The skin panels in the nose region, engine ramps and nozzle region, and the combustion module are one side of a heat exchanger system that pumps aerodynamic heating into an energy extraction loop. The very cold hydrogen passes through a skin panel that absorbs the incoming aerodynamic heating. The energy extraction loop lowers the hydrogen temperature and then passes it to another heat exchanger panel. Thus the liquid hydrogen goes through a series of net energy additions until it reaches the combustion chamber where is injected as a hot gas (Figure 4.18). This concept goes back to the original aerospace plane for the United States Air Force to which The Marquardt Company was a contractor. At that time John Ahern worked with Charles Lindley, Carl Builder and Artur Magar, who originated many of these concepts.

Figure 4.18 depicts a typical closed-loop heat pump loop identified in Figure 4.17 as a rectangle with "EX" inside, and the fuel wall injection system. This particular loop is for one of the inlet ramps ahead of the engine module. The three heat exchangers form a closed-loop system where thermal energy extracted from the skin panels is used to power an expansion turbine that drives the working fluid compressor. The net work exacted can be used to power electrical generators, hydraulic pumps, refrigeration units or fuel boost pumps. With hydrogen as fuel, the vehicle is independent of ground power sources and can self-start as long as there is hydrogen in the fuel tanks. Eventually the fuel reaches the engine module where it picks up the heat transferred to the combustor walls. When the hydrogen reaches its maximum temperature it is injected into the combustion chamber via series of Mach 3 nozzles at a low angle to the wall. The size of the nozzles can be small and approach the equivalent of a porous wall. The result is that the hydrogen acts as film cooling for the wall, reducing the wall friction and heat transfer rate. For a Mach 3 wall nozzle the kinetic energy of the injected fuel also creates thrust, and the thrust per unit fuel flow, 7sp, is given in equations (4.10) for hydrogen.

Fuel Isp = 9.803 T0:5197 (s) T is in Rankine Fuel Isp = 13.305 T0 5197 (s) T is in Kelvins

3,000 r

3,000 r

Builder's analysis No drag losses

15 20

Flight speed (kft/s)

Figure 4.19. System thermal integrated specific impulse.

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