This paper presents algorithms to calculate supersonic flow about a prospective ring wing launch vehicle by the marching method and the relaxation method. The feature of the algorithms is the introduction of two computational subregions in the ring wing zone over the rocket airframe. In the marching algorithm, the computation region is reconstructed according to the position of the marching cross-section relative to the leading and trailing edge of the ring wing. When it finds itself at the leading edge of the ring wing, the computational region is split into a lower subregion between the rocket airframe and the downstream face of the ring wing and an upper subregion between the upstream face of the ring wing and the bow shock front. When the marching cross-section finds itself at the trailing edge of the ring wing, the lower and the upper computational subregions are merged into a single computational region. Based on the marching algorithm and using the authors’ rocket flow calculation program, software is developed for a fast numerical calculation of supersonic flow about ring wing rockets. For a particular ring wing rocket configuration, the paper presents the results of comparative calculations of supersonic flow about the rocket in the form of gas-dynamic parameter isolines in the flow field and the pressure distribution over the rocket airframe and the ring wing. The results for the marching method and the relaxation method are compared. It is shown that the ring wing is responsible for an undulatory pressure distribution between the rocket airframe and the downstream face of the ring wing. The marching method simulates the flow pattern between the rocket airframe and the downstream face of the ring wing more adequately, and its computation time is two orders of magnitude shorter than that of the relaxation method. The relaxation method should be used in the case of subsonic flows between the rocket airframe and the downstream face of the ring wing. The algorithm and software developed are recommended for parametric calculations of supersonic flow about ring wing rockets.

This paper discusses the use of the authors’ fast methods and programs for the calculation of 3D supersonic flow about a flying vehicle and thermogas dynamic processes in the components of an airframe-integrated ramjet. To conduct fast comprehensive calculations, use is made of marching methods, which are two to three orders of magnitude faster than pseudoviscosity methods. 3D supersonic flows about the airframe, in the inlet section of the air intake, and in the exhaust jet are calculated using a “viscous layer” model or Godunov’s scheme for the inviscid approximation. Subsonic flows in the outlet section of the air intake and in the combustion chamber are calculated using a “narrow channel” or a quasi-one-dimensional model. The elements of the presented methods and programs that complement a previously proposed fast comprehensive model are described in more detail. A method for assigning the spatial shape of the flying vehicle surface and the ramjet duct walls is described. A simplified approach to determining the critical area of the exit nozzle in the one-dimensional approximation is proposed. The paper substantiates the advantages of marching methods over pseudoviscosity ones in the predesigning of ramjets with direct account for flow choking, which may occur in the combustion chamber or the exit nozzle. The calculated 3D flows in the individual components and the full assembly of a stylized-shape flying vehicle are presented. The main advantages of the proposed methods and programs are their comprehensiveness and fast computation speed. Their use in the calculation of 3D supersonic flow about a ramjet flying vehicle shortens the ramjet component predesigning time.