A new generation of spacecraft is currently being designed. Some of the proposed concepts involve the use of air breathing engines during part of the earth to orbit flight phase. In the case of the HOTOL concept studies, the engine intakes will be covered for the re‐entry phase, and will protrude through the windward surface shock layer during re‐entry. An understanding of the complex flow which will occur over the closed intakes during the hypersonic re‐entry is important for at least two reasons. Firstly, the heat transfer on the surfaces has to be estimated to allow for suitable intake cover design. Secondly, the wake of the intakes interacts with the underside of the wings and control surfaces, and could possibly cause handling anomalies.The present paper describes a study in which a simplified model involving a double wedge mounted on a flat plate at incidence (Fig. 1) was tested in the Free Piston Shock Tunnel T3 at the Australian National University. Heat transfer measurements and shock luminosity photographs were recorded at two operating conditions, one with a stagnation enthalpy of 22 MJ/kg and the other with 2.8 MJ/kg. A flow analysis which identified a number of significantly different flow regimes was also performed.Heat transfer measurements indicate that heating rates well in excess of those expected at the stagnation point on the nose of the spacecraft can be expected. The results also highlighted a compromise which is a necessary feature of this type of design. The compromise involves a trade off between intake efficiency during the air breathing phase of operation and the reduction of heat transfer during the re‐entry phase.