Description: Mars sample return ascent vehicles require a velocity change exceeding 4 km/s within minutes of launch and higher propellant mass fractions than are offered by current pressure-fed satellite propulsion systems. Pump-fed systems are potentially low-mass, but a suitable high-pressure nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) bipropellant rocket engine remains to be developed. The state of the art in NTO/MMH combustion chambers is iridium-lined rhenium (Ir/Re) manufactured by Ultramet and used by Aerojet in its HiPAT engine. In previous and current work, Ultramet has demonstrated the potential for fabrication of refractory metal and ceramic matrix composite (CMC)-lined carbon/carbon (C/C) components, thereby combining the benefits of the refractory liner with the substantially reduced weight and cost of C/C. The shortfall of using C/C as the lightweight structure is the disparity in thermal expansion, causing highly stressed components. In this project, Ultramet will demonstrate a processing route to integrally bond liners of a zirconium carbide (ZrC) CMC and of rhenium to C/C that will be applied to the fabrication of a Mars sample return ascent engine combustion chamber. Phase II will culminate in the design and fabrication of a 1000-N, 500-psi refractory-lined C/C chamber that will be delivered to NASA. To facilitate testing within project funding resources, a 22-N 500-psi refractory-lined C/C chamber will also be designed, fabricated, and hot-fire tested at one of the major propulsion system integrators (Aerojet, Moog-ISP, or Northrop Grumman), all of which have expressed interest in teaming on the project.

Benefits: Successful development and demonstration of CMC- and rhenium-lined carbon/carbon for advanced liquid rocket engine components will result in a streamlined manufacturing technique that costs less and yields lighter components than current techniques. Less rhenium would be needed for a given strength, which would decrease propulsion system mass and cost. The increase in allowable operating temperature limit would improve performance and capability on future satellites and enable new missions. Application to advanced combustion chambers will result in less expensive, lighter, and more capable components than are available currently. The proposed material can be used in any NASA systems that rely on rhenium, such as NTO/MMH bipropellant iridium/rhenium apogee kick motors and advanced green ionic liquid monopropellant (HAN- and ADN-based) thrusters for satellites and planetary probe in-space propulsion and planetary and other large body ascent vehicles. Other applications include orbit transfer, maneuvering, station keeping, and attitude control for satellites.

The proposed refractory-lined C/C chambers can be used in attitude control and apogee engines for commercial and government satellites and solid and liquid divert and attitude control system components for missile interceptor kinetic kill vehicles. Other potential applications include gas generators for auxiliary power units and tank pressurization systems, and liquid gun propellant systems. The proposed concepts will benefit NASA, DoD, and commercial rocket propulsion systems that rely on rhenium, and extensions of the technology can be used in applications such as heat pipes.