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.

The demand of small payload launch vehicles has been growing over the past years. Customers base their selection of launch vehicles on cost-effectiveness, flexibility, availability and reliability. A new launch vehicle, the Dawn Aerospace Mk-III, is proposed to be developed, while designed to take into account all these criteria.

For flexible and frequently available operations the launch trajectory is integrated in the common airspace. Manoeuvrability is identified as a key capability for safe operations in the common airspace. For high manoeuvrability the first stage is designed as a rocket propelled airplane, a so-called 'spaceplane'. The expendable upper stage is stored internally. After payload injection the first stage returns to the spaceport of take-off. This means the System is a Two Stage to Orbit semi-Reusable Launch Vehicle integrated in the common airspace.

Cost-effectiveness is a primary selection criterium in the decision making of customers, which is why cost is included from an early stage in the development. This study shows that for identifying the cost gradient in the design space the total dry mass of the vehicle is sufficient. In this way cost optimality is included, although the Cost per Flight cannot be determined. What is determined is the effect of different technical and operational considerations. Taking into account qualitative cost differences, a selection of cost derived mission scenarios are studied. This includes different Return to Launch Site methods, first stage engine design and lay-out, the prohibition of fairing usage and integrated landing gear for take-off and landing.

To analyse and optimize the different designs a Multidisciplinary Design Optimization tool is developed. This tool optimizes the vehicle and the ascent trajectory simultaneously to determine the lowest total dry mass solution meeting all requirements and constraints. To estimate the aerodynamic performance of the first stage the X-34 Advanced Technology Demonstrator is used as a reference vehicle. This means the geometry of the first stage is not optimized, while the size is. The upper stage is modelled as a conventional upper stage, of which the size and geometry is optimized.

The study shows that the proposed design is feasible, meeting all requirements and constraints. The result is a vehicle with a total dry mass of 6273.0 kg, Gross Take-Off Weight of 42972.8 kg and a first stage length of 19.4 m. Of the total dry mass 94.5% is reusable. The return of the first stage is driving the trajectory design, as a steep ascent trajectory is required for limiting the downrange of the first stage. This results in 30%-50% more gravity loss in the System, which demands for a high ΔV performance. Due to the size of the first stage and the propellant required for returning, the first stage ascent ΔV is limited. For that reason, the upper stage design has a Propellant Mass Fraction of 0.939, increasing the upper stage ascent ΔV performance. Achieving such a Propellant Mass Fraction is possible, but challenging. The upper stage design is identified as a key element in the System performance for successfully meeting all mission requirements.

Three different Return to Launch Site methods are compared. Two methods, in-plane pitch over and aeroturn, are active which requires return propellant. The third method, glideback, is passive. The study shows that glideback can be favourable. The total dry mass increase is 4.4% when compared to in-plane pitch over. The increase in heat load is ~16%, but the heat load in this study is limited with a total heat load of ~2.0 MJ/m^2. However, for glideback an even higher upper stage Propellant Mass Fraction of 0.946 is required. This means that the result of the upper stage design determines if this return method is feasible.

Using a shared engine design on the first and upper stage shows promising results. Using a single first stage engine reduces the first stage dry mass by 4.0%. However, using a shared engine design is expected to decrease the development cost drastically, due to the reduced size of individual engines. The prohibition of fairing usage shows a stronger effect on the result. Allowing the use of a fairing decreases the first stage dry mass by a maximum of 9.6%. Fairing usage on the other hand harms the safe operations in the common airspace as the ejection of uncontrolled material requires large safety zones. The penalty on mass is acceptable for allowing integration in the common airspace. A landing gear sized for take-off results in a heavier vehicle. When the first stage is supported by a cart during take-off the first stage dry mass is decreased by 24.6%. When the first stage is air-launched the first stage dry mass decreases by 33.3%.