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.

VX-200 Magnetoplasma Thruster Performance Results Exceeding Fifty-Percent Thruster Efficiency

Abstract: H IGH-POWER electric propulsion thrusters can reduce propellant mass for heavy-payload orbit-raising missions and cargo missions to the moon and near-Earth asteroids, and they can reduce the trip time of robotic and piloted planetary missions [1–4]. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200 engine is an electric propulsion system capable of processing power densities on the order of 6 MW=m with a high specific impulse (4000 to 6000 s) and an inherent capability to vary the thrust and specific impulse at a constant power. The potential for a long lifetime is due primarily to the radial magnetic confinement of both ions and electrons in a quasi-neutral flowing plasma stream, which acts to significantly reduce the plasma impingement on the walls of the rocket core. High-temperature ceramic plasma-facing surfaces handle the thermal radiation: the principal heat transfer mechanism from the discharge. The rocket uses a helicon plasma source [5,6] for efficient plasma production in the first stage. This plasma is energized further by an ion cyclotron heating (ICH) RF stage that uses left-hand polarized slow-mode waves launched from the high field side of the ion cyclotron resonance. Useful thrust is produced as the plasma accelerates in an expanding magnetic field: a process described by conservation of the first adiabatic invariant as the magnetic field strength decreases in the exhaust region of the VASIMR [7–9]. End-to-end testing of the VX-200 engine has been undertaken with an optimum magnetic field and in a vacuum facility with sufficient volume and pumping to permit exhaust plumemeasurements at low background pressures. Experimental results are presented with the VX-200 engine installed in a 150 m vacuum chamber with an operating pressure below 1 10 2 Pa (1 10 4 torr), and with an exhaust plume diagnostic measurement range of 5 m in the axial direction and 1 m in the radial directions. Measurements of plasma flux, RF power, and neutral argon gas flow rate, combined with knowledge of the kinetic energy of the ions leaving the VX-200 engine, are used to determine the ionization cost of the argon plasma. A plasmamomentum flux sensor (PMFS)measures the force density as a function of radial and axial positions in the exhaust plume. New experimental data on ionization cost, exhaust plume expansion angle, thruster efficiency, and total force are presented that characterize the VX-200 engine performance above 100 kW. A semiempirical model of the thruster efficiency as a function of specific impulse has been developed to fit the experimental data, and an extrapolation to 200 kWdc input power yields a thruster efficiency of 61% at a specific impulse of 4800 s.

Compressible Wall-Injection Flows in Laminar, Transitional, and Turbulent Regimes: Numerical Prediction

Abstract: Numerical simulations of compressible rocket flows are conducted in laminar, transitional, and turbulent regimes. The laminar simulation is carried out on a planar rocket flow without nozzle using the unsteady two-dimensional Navier-Stokes system. The transitional and turbulent flows are performed in three-dimensional on an extended rocket geometry with a divergent outlet using compressible large eddy simulation (LES) models. In both cases, the compressibility effect plays an important role. In the laminar case, pressure oscillation is forced at the outflow boundary. The time-averaged part of the solution is compared with the inviscid theory of compressible rocket flow of Balakrishnan et al. (Balakrishnan, G., Linan, A., and Williams F. A., "Compressibility Effects in Thin Channels with Injection," AIAA Journal, Vol.29, No. 12, 1991, pp. 2149-2154) and the oscillatory part with the acoustic layer model of Majdalani and Van Moorhem (Majdalani, J., and Van Moorhem, W. K., "Improved Time-Dependent Flowfield Solution for Solid Rocket Motors," AIAA Journal, Vol. 36, No. 2, 1998, pp. 241-248). The mean flow from the present numerical result is in better agreement with the compressible theory than the conventional Taylor's profiles (Taylor, G. I., "Fluid Flow in Regions Bounded by Porous Surfaces," Proceedings of the Royal Society of London, Series A: Mathematical and Physical Sciences, Vol. 234, 1956, pp. 456-475), as expected. The oscillatory part of the flow agrees well in the first quarter of the axial extent, near the head end. Farther downstream, the discrepancies develop rapidly between the numerical result and the acoustic-layer model. Possible causes of the difference are the effect of compressibility, which alters the local speed of sound, hence, acoustic properties, and the interference of hydrodynamic instabilities. In the transition and turbulent regimes, the dynamic LES model is applied on different resolutions. The measurements data of Traineau et al. (Traineau, J. C., Hervat, P., and Kuentzmann, P., "Cold Flow Simulation of a Two Dimensional Nozzleless Solid Rocket Motor," AIAA Paper 86-1447, June 1986) are employed for comparison purposes. The refinement study by comparison with the measurement data suggests the importance of resolving the laminar and transition region for a reliable application of LES in transitional flows. With the consideration of this aspect, LES with efficient grid size can produce resonable accuracy. Forcing hydrodynamic instabilities and a more realistic injection fluctuations model are recommended.

Solid rocket exhaust in the stratosphere - Plume diffusion and chemical reactions

Abstract: A model has been developed to examine, on a local scale, the reactions of rocket exhaust from solid rocket motors with stratospheric ozone. The effects were examined at two different altitudes. Results of the modeling study indicate that afterburning chemistry of reactive exhaust products can cause local but transient (on the order of several minutes) loss of ozone. The modeling study included potential heterogeneous reactions at aluminum oxide surfaces. Results indicate that these potential heterogeneous reactions do not have a major impact on the local plume chemistry. Homogeneous reactions appear to be of more consequence during the early dispersion of the plume. It has also been found that the rate of plume dispersion has a very significant effect on local ozone loss.

Nuclear Thermal Rocket/vehicle design options for future NASA missions to the Moon and Mars

Abstract: The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a reference 240 t-class heavy lift launch vehicle (HLLV) and smaller 120 t HLLV option. Attractive performance characteristics and high-leverage technologies associated with both the engine and stage are identified, and supporting parametric sensitivity data is provided. The potential for commonality of engine and stage components to satisfy a broad range of lunar and Mars missions is also discussed.