Topic
Rocket
About: Rocket is a research topic. Over the lifetime, 14018 publications have been published within this topic receiving 95852 citations. The topic is also known as: rockets.
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TL;DR: The data from the third rocket of the Substorm-GEOS series have been studied as discussed by the authors, which was launched from ESRANGE on January 27, 1979 at 21.20 UT into a pulsating aurora.
Abstract: Data from the third rocket of the Substorm-GEOS series have been studied. The rocket was launched from ESRANGE on January 27, 1979 at 21.52.20 UT into a pulsating aurora.
104 citations
TL;DR: In this article, a number of studies have been directed toward delineation of the mechanism and magnitude of nonequilibrium aspects of two-phase nozzle expansion processes, including the effects of thermal and velocity lags.
Abstract: T USE of metallic fuel constituents in modern rocket engines has brought attention to nonequilibrium aspects of two-phase nozzle expansion processes. Since condensed metal oxide combustion products (which comprise 30 to 40% by weight of the total products of contemporary solid rockets) can do no expansion work, their presence in the rocket nozzle can only be deleterious to the effectiveness of the nozzle expansion process in converting thermal to kinetic energy. The condensed particles are accelerated in a nozzle almost exclusively by drag forces associated with lag or slippage of the particles relative to the expanding gas. Some performance loss relative to the calculated ideal no-slip expansion process must always be associated with macroscopic size particles, and experience has shown that the magnitude of the loss increases with the weight fraction of particles. Significant velocity and thermal lags thus have been suspected as a prime cause of rocket performance losses, and a number of studies have been directed toward delineation of the mechanism and magnitude of these lag effects. Early studies were summarized and extended in the review of Altman and Carter (l). Primarily, these early studies served to place bounds on the performance losses by examining the limiting cases of no-lag and complete lag. They demonstrated that thermal lag ordinarily has a lesser effect on specific impulse than does velocit}^ lag. Gilbert, Davis, and Altman (2) were the first to relate the losses to particle size. They solved the linear equation that results from assuming the drag force to be proportional to the velocity difference (Stokes' law) for the case of linearly accelerated nozzle gas. They demonstrated that, typically, a 1-ju diam particle follows the gas velocity closely, whereas a IQ-fj, diam particle has a significant lag. All of these early studies treated the nozzle expansion processes as though they are uncoupled, i.e., the thermal lag
103 citations
24 Aug 1964
103 citations
TL;DR: In this paper, the transition from free shock separation to restricted shock separation in a parabolic nozzles was analyzed and the cap-shock pattern was identified to be the cause of this transition.
Abstract: Uncontrolled flow separation in nozzles of rocket engines is not desired because it can lead to dangerous lateral forces. Different origins for side loads were identified in the past. Meanwhile, it is proven that in thrust-optimized or parabolic nozzles, a major side load occurs as a result of the transition of separation pattern from free shock separation to restricted shock separation and vice versa. Reasons for the transition between the separation patterns are discussed, and the cap-shock pattern, which is identified to be the cause of this transition, is closely analyzed. It turns out that this pattern can be interpreted as an inverse Mach reflection of the internal shock at the nozzle axis. To prove the transition effect as main side-load driver, a subscale test campaign has been performed. Two different nozzle contours, a thrust-optimized and a truncated ideal nozzle with equal performance data, were tested. Highest side loads were measured in the thrust-optimized nozzle, when the separation pattern changes from free to restricted shock separation. Side loads measured in the truncated ideal nozzle were only about one-third as high as in the thrust-optimized nozzle.
102 citations
TL;DR: In this paper, the performance of a single-stage-to-orbit aerospace plane with a fixed-geometry combined-cycle engine was analyzed with a simple simulation model and the cooling requirement of the engine and the pitching moment of the plane were investigated.
Abstract: Operating conditions and performances of a fixed-geometry combined-cycle engine for a single-stage-to-orbit aerospace plane were calculated with a simple simulation model. With the flow conditions calculated with the model, the cooling requirement of the engine and pitching moment of the plane were investigated. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. The engine had a fixed geometry in its operation. Subsonic combustion was conducted with no second throat in the combustor under the ejector-jet mode and the ramjet mode. Propellants were liquid hydrogen and liquid oxygen. The coolant flow rate became larger than the fuel flow rate. The excessive flow rate decreased the specific impulse above Mach 9 and restricted application of the airbreathing engine mode up to Mach 11. The pitching moment of the plane would be balanced even in the space in the configuration with the combined-cycle engine mounted on the windward surface.
102 citations