One of the main issues of the international space programs is the identification and the design of a more flexible and possibly reusable manned space vehicle, able to re-enter from space missions and safely land on earth. Several re-entry strategies are feasible -from classical, direct capsule re-entry (APOLLO, SOYUZ) over a skip re-entry (ZOND, ORION) to aerocapture/aerobraking with a following orbital re-entry (Space Shuttle). But to achieve the objectives mentioned above, vehicles with a high lift-over-drag (L/D) ratio are necessary. A concept for this high L/D earth re-entry vehicle is the PHOEBUS (Plane-shaped Hypersonic Orbital Re-entry BUS) of the OHB-Team. This concept of a high L/D vehicle was studied within a framework of an ESA study and performs long duration re-entry trajectories with large down-and crossrange and good maneuverability by proper utilization of the lift forces. This paper presents an overview of the PHOEBUS vehicle design which is based on the OHB SpacePlane LR-L. Here the characteristic of a high L/D vehicle with a hot structure concept is described, using state-of-the-Art materials and an innovative design to permit low-risk re-entry flights with lower aerothermal -and deceleration loads. Furthermore, the flexibility to fly very large re-entry trajectories by the use of innovative entry strategies -as well as the limits for such strategies is described.A first iteration of the system performance related to the aerodynamic -and aerothermodynamic aspects as well as the downrange -and crossrange capabilities of the PHOEBUS vehicle are shown. The significantly larger performance of PHOEBUS will be illustrated by comparison of these parameters with conventional re-entry vehicles.

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A method for the evaluation of launch vehicle performance is presented. The method avoids the CPU-intensive computations needed to nd a launch trajectory by a numerical optimization program, and it is based on variation of velocity balance formulas previously introduced. These formulas have been extended to allow a more realistic model of the launcher, including multistaging, variable thrust, nonlinear mass rate, etc. The results obtained show that accuracy is good, because the error in the estimate of the maximum payload mass that can be inserted into required orbits is below 10%. Moreover, the formulas developed can provide, in a very short time, a complete trajectory that can be a suitable input for trajectory optimizers.

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