Aerothermal Analysis of Small Payloads Delivered Into Low Earth Orbit From an Airborne Launch Platform

In recent years there has been an ever increasing need to launch small payloads (∼1–100 kg) into low earth orbit (LEO). Examples include the defense and telecommunications industries. Permanent human presence in LEO, such as the international space station, requires continual re-supply from earth. Additionally, NASA’s stated mission of launching a manned mission to Mars requires many tonnes of raw materials to be economically launched into LEO and assembled there. Conventional rocket launch from earth is prohibitively expensive for small mass payloads. Estimates range from $7000–$20,000 to launch 1 kg of mass into low earth orbit. Several concepts have been proposed to economically launch small payloads from earth, including light gas guns, electromagnetic launchers and the so called “slingatron” concept. The goal of these concepts is to reduce the cost per kg (to under $1000) to achieve LEO. Each of these concepts involves launching small payloads that traverse the atmosphere and then placed into low earth orbit using thrusters to turn the velocity vector into a stable circular orbit. As the launch vehicle traverses the dense lower portion of the atmosphere it experiences severe thermal heating loads that must be absorbed by a thermal protection system (TPS) if the payload is to survive the transit. The University of Texas is currently heading a multi-university research initiative (MURI) to study the feasibility of launching small payloads into low earth orbit from an electromagnetic gun housed in an airborne platform. As part of the study, the aerothermal issues associated with traverse through the atmosphere and propellant mass required to achieve a stable circular orbit are investigated. The effort focuses on quantifying the required parasitic mass of the thermal protection system (TPS) and propellant need to place a nominal 10 kg launch mass into a circular low earth orbit from an electromagnetic launcher at 16 km altitude. The TPS is assumed to be graphite or carbon-carbon composite. In this effort, we consider ballistic trajectories only. Circular orbit is achieved using rocket thrusters at the terminal altitude. Total parasitic mass (TPS + propellant) is estimated for various launch angles.