The aim of this work is to study the effect of disturbances from startups and shutdowns of a restartable sustainer engine on the operation of control jet engines in their continuous and pulse operation with account for the integration of the engine feed lines. Abandoning the traditional feed of control engines from individual tanks increases the payload mass by eliminating the need for a gas displacement system and offers a more complete use of the onboard propellant. The main objectives of the system of control engines as actuators of the launch vehicle upper stage flight control system are roll, pitch, and yaw control in different operating regimes of the sustainer engine, the acceleration of the stage before sustainer engine restarts, and the removal of the stage to an utilization orbit with remaining propellant burn-up. Because the feel lines of the sustainer engine and the control engines are hydraulically connected, at sustainer engine startups/shutdowns the propellant component pressure at the control engine inlets is subject to disturbances in the form of surges and dips, in response to which the rocket stage flight control system must generate additional control actions. The control engine operation under disturbances from the sustainer engine was studied using the authors' comprehensive mathematical model, which describes the propellant component flow in the feed lines, electrically operated fuel valves, and combustion chambers of control engines, and the time profiles of disturbances recorded in full-scale ground tests of the Cyclon-4M launch vehicle upper stage. Calculations were conducted for the most strenuous combinations of the control engine operation under disturbances from sustainer engine startups and shutdowns. The calculated data show that the control engine thrust is within the limits specified in the requirements specification for the development of the control engines as a part of the liquid-propellant jet system.
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.
The aim of this work is to determine the hydraulic resistance of a gas-dispersed flow in a horizontal channel with a minimum of empirical data. Assuming the gas-dispersed flow energy balance as a sum of two terms: the carrier gas energy and the particle energy, a relation for hydraulic resistance determination in Gasterstadt's form is obtained. To find the particle drag and relative velocity appearing in this relation, use is made of the results of a numerical solution of the problem of the motion of particles in a horizontal channel in Lagrangian variables with account for the interaction of the particles with the channel walls and with one another. The effect of the particles on the carrier gas parameters is neglected. A verification of the proposed technique using various empirical relationships reported in the literature has shown that the calculated and the experimental results are in satisfactory agreement. The proposed technique may be used in the development and design of engineering systems with gasdispersed flows.
Air intake shape designing is the key problem in the development of a ramjet. The aim of this paper is to formulate an algorithm for on-the-fly computing of supersonic flow stagnation in the passage of an air intake at its predesigning stage. Consideration is given to computing flows in the passages of counter-pressure air intakes by the marching method using a quasi-one-dimensional approach to computing the subsonic flow at the passage outlet and by the time relaxation method. The efficiency of these methods is compared. It is suggested that the gas-dynamic flow parameters be determined at the predesigning stage using the algorithm of on-the-fly computing by the marching method with the determination of the normal shock position for which the required flow velocity coefficient at the outlet section of the air intake passage is realized.
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