The commercialization of the space industry has led to a reduction in size and weight of satellites and launching vehicles, which have effectively reduced the cost of space services. The costs of launching payloads to space can be significantly reduced when the launching vehicles are reusable. The two-stage-to-orbit system with a winged, reusable first stage vehicle, is deemed to offer benefits in terms of operations, as it can take-off and land from a runway.
Dawn Aerospace is pursuing development of the winged semi-Reusable Launching Vehicle with horizontal take-off and landing capabilities in the Mk-III concept. In the previous work on this concept, the aerodynamic design has not been addressed. It is important that the aerodynamic design is already considered in the conceptual design, as the large variations in flight conditions during the mission pose conflicting requirements on the aerodynamic design.
A performance analysis model of the launching vehicle was developed, in which the aerodynamic discipline is integrated. Using the developed model, sensitivity analyses and case studies were performed to investigate the impact of design changes on the mission performance. These results indicate that the limited gliding range of the first stage vehicle influences the propellant mass fraction of the upper stage.
The performed case studies indicate how the propellant mass fraction of the upper stage can be influenced, by changes to the vehicle design. Analysis of an alternative wing concept that improves the gliding performance of the first stage vehicle shows that the propellant mass fraction of the upper stage vehicle can be reduced from 91.5% to 90.3%. This can be achieved as the improved gliding range allows to reduce the delta-V delivered by the upper stage, and a less steep ascent trajectory that results in 8.9% less gravity losses. Analysis of a changed fuselage configuration indicates a reduction of propellant mass fraction from 91.5% to 90.4%.
Also changes to the trajectory design were analyzed, and by either reversing the direction of the take-off maneuver or by reserving propellant for the return trajectory, the propellant mass fraction can be reduced. The reductions achieved in the case studies were 0.5% for the reversed take-off direction, and 1.2% for the propelled return trajectory.