Description: The specific impulse of a rocket engine increases as the chamber pressure increases, but so does the heat flux to the chamber wall. Ultimately, this defines the maximum operating pressure for the engine. For regeneratively cooled engines, even those using film cooling, the practical limit has been reached, and further increases in chamber pressure are simply not possible. Transpiration cooling does not have this limitation. Furthermore, because a transpiration-cooled engine pumps only a tiny fraction of the fuel through the wall, a smaller and hence lighter pump can be used, which will significantly reduce the dry mass. Finally, because transpiration cooling can keep the wall much cooler than regenerative cooling with film cooling, a transpiration-cooled engine can use less refractory (i.e., lighter weight) materials, thereby achieving additional reductions in dry mass. The net results are significant increases in the thrust-to-weight ratio and specific impulse and a significant decrease in the dry mass of the system. The perceived limitation of transpiration cooling with a porous wall is coking and blockage of the pores if a carbon-based fuel such as methane is used. In previous work using LOX/H2 propellant, Ultramet showed that with minimal transpiration flow, the wall temperature can be kept well below the point at which methane would form coke. In this project, Ultramet will work with Purdue University to build on previous success with transpiration cooling in LOX/H2 engines and design a lightweight LOX/LCH4 engine in the 10,000- to 25,000-lbf thrust range. The transpiration model will be physics-based and applicable to both LOX/LCH4 and LOX/H2. Key component demonstrators will be fabricated and used to collect empirical data on the thermal, structural, and hydraulic characteristics of the wall architecture. Transpiration rates on subscale hardware will be verified through flow testing, and empirical data will be used to verify the predicted lack of coking.

Benefits: The primary NASA applications will be for high-performance, high-thrust O2/CH4 rocket engines in the 10,000- to 25,000-lbf thrust class, which can be used on booster upper stages, as well as for ascent/descent engines for Mars and lunar landers. Smaller engines in the 1000-lbf thrust class can be used for Earth departure stages and orbital maneuvers of large spacecraft. Larger engines in the 500,000-lbf thrust class can be used for lower stages on heavy-lift launch vehicles. This technology can be applied to LOX/H2, LOX/LCH4, and LOX/RP-1 engines.

Non-NASA applications for this technology include upper stage booster engines, main engines for smaller commercial or military launch vehicles, and main engines on air-launched vehicles for delivering payloads to low Earth orbit. Larger engines in the 500,000-lbf thrust class can be used for lower stages on commercial and military heavy-lift vehicles. Because the technology is not propellant-specific, it can also be applied to LOX/LH2 and LOX/RP-1 engines in addition to LOX/LCH4.