Showing papers on "Liquid-propellant rocket published in 1986"

Comprehensive modeling of a liquid rocket combustion chamber

Abstract: An analytical model for the simulation of detailed three-phase combustion flows inside a liquid rocket combustion chamber is presented. The three phases involved are: a multispecies gaseous phase, an incompressible liquid phase, and a particulate droplet phase. The gas and liquid phases are continuum described in an Eulerian fashion. A two-phase solution capability for these continuum media is obtained through a marriage of the Implicit Continuous Eulerian (ICE) technique and the fractional Volume of Fluid (VOF) free surface description method. On the other hand, the particulate phase is given a discrete treatment and described in a Lagrangian fashion. All three phases are hence treated rigorously. Semi-empirical physical models are used to describe all interphase coupling terms as well as the chemistry among gaseous components. Sample calculations using the model are given. The results show promising application to truly comprehensive modeling of complex liquid-fueled engine systems.

Tripropellant rocket engine

Abstract: The invention combines liquid hydrogen with liquid oxygen and a hydrocarbon as the fuels in a tripropellant rocket booster engine.

Experimental thrust performance of a high-area-ratio rocket nozzle

Abstract: An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

Method for the operation of a liquid-fuelled rocket engine and rocket engine for implementing the method

Abstract: Method for the operation of a liquid-fuelled rocket engine, comprising a combustion chamber with propelling nozzle, propellant pumps, one or more pump drive turbines, which are driven by one of the two propellants, especially hydrogen, which has previously flowed through the hot propelling nozzle and combustion chamber wall, heat from a heat exchanger (11) being imparted to the previously heated propellant, especially the hydrogen before it enters the turbine or turbines (8 and 9) as turbine propellant gas (Ht), to which heat exchanger combustion gases (B) are admitted which are produced from partial amounts of the rocket propellants in a stoichiometrically driven auxiliary combustion chamber (10), the exhaust gases (AB) from which are introduced into the propelling nozzle (3) (Fig. 1). … …

Bipropellant rocket exhaust plume analysis on the Galileo spacecraft

Abstract: This paper describes efforts to quantify the contaminant flow field produced by 10 N thrust bipropellant rocket engines used on the Galileo spacecraft. The prediction of the composition of the rocket exhaust by conventional techniques is found to be inadequate to explain experimental observations of contaminant deposition on moderately cold (200 K) surfaces. It is hypothesized that low volatility contaminants are formed by chemical reactions which occur on the surfaces. The flow field calculations performed using the direct simulation Monte Carlo method give the expected result that the use of line-of-sight plume shields may have very little effect on the flux of vapor phase contaminant species to a surface, especially if the plume shields are located so close to the engine that the interaction of the plume with the shield is in the transition flow regime. It is shown that significant variations in the exhaust plume composition caused by nonequilibrium effects in the flow field lead to very low concentrations of species which have high molecular weights in the more rarefied regions of the flow field. Recommendations for the design of spacecraft plume shields and further work are made.

Enhanced heat transfer combustor technology, subtasks 1 and 2, tast C.1

Abstract: Analytical and experimental studies are being conducted for NASA to evaluate means of increasing the heat extraction capability and service life of a liquid rocket combustor. This effort is being conducted in conjunction with other tasks to develop technologies for an advanced, expander cycle, oxygen/hydrogen engine planned for upper stage propulsion applications. Increased heat extraction, needed to raise available turbine drive energy for higher chamber pressure, is derived from combustion chamber hot gas wall ribs that increase the heat transfer surface area. Life improvement is obtained through channel designs that enhance cooling and maintain the wall temperature at an accepatable level. Laboratory test programs were conducted to evaluate the heat transfer characteristics of hot gas rib and coolant channel geometries selected through an analytical screening process. Detailed velocity profile maps, previously unavailable for rib and channel geometries, were obtained for the candidate designs using a cold flow laser velocimeter facility. Boundary layer behavior and heat transfer characteristics were determined from the velocity maps. Rib results were substantiated by hot air calorimeter testing. The flow data were analytically scaled to hot fire conditions and the results used to select two rib and three enhanced coolant channel configurations for further evaluation.

Comparison of methane and propane rockets

Abstract: Martin (1983) compared methane and propane fuels for a single-stage-to-orbit vehicle, demonstrating a significant advantage for propane. Attention is presently given to the ways that this comparison changes when both methane and propane vehicles are optimized. It is found that while the difference is slightly reduced, propane remains the better hydrocarbon fuel for dry mass minimization.

JANNAF (Joint Army-Navy-NASA-Air Force) Propulsion Meeting (1986) Held in New Orleans, Louisiana on 26-28 August 1986. Volume 1.

Abstract: : This volume, the first of five volumes, contains 54 unclassified, unlimited distribution papers that were presented at the 1986 JANNAF Propulsion Meeting. Specific subjects discussed include asbestos-free insulation, production engineering, solar thermal propulsion, stress corrosion of metals, combustion modelling in airbreathing engines and ramjets, chemical analysis of liquid gun propellants, production and processing of gun propellants, rocket motor nozzle structural analysis and instrumentation, electrothermal thrusters, studies of hydrazine compatibility and hazards, and hydrogen/oxygen propellants for space applications. Keywords: Electric propulsion; Liquid propellant rocket engines; Rocket propulsion; Solid rocket propellants; Space propulsion.

Future LOX/Hydrocarbon Booster Engines

Abstract: This paper discusses several LOX/Hydrocarbon booster engines that are being considered for future launch vehicles. The various concepts are compared. Introducing liquid hydrogen to the liquid oxygen/hydrocarbon booster engine appears to offer many benefits.

Rocket plume impingement heat transfer on plane surfaces

Abstract: An analytic procedure has been derived for the rapid calculation of maximum heat flux on a flat plate due to the impingement of liquid rocket exhaust plumes. The impingement configurations considered generally lead to plume flow in the transitional flow regime, described in the approximation of a source flow expression. Newtonian impact analysis is employed to determine the location of the maximum pressure, considered to result in a stagnation point within the Newtonian shock layer. The Detra and Hidalgo atmospheric re-entry heat-transfer equation is modified to account for the effects of oblique plume flow incidence and related asymmetric shocklayer flow in the neighborhood of the stagnation point. The modified re-entry heating expression, including a single correlation factor of the order of unity, is shown to agree with extensive test data available on liquidbipropellant heat transfer reported by Piesik et al. The analytical procedure yields improved correlation of the peak heat-transfer data and, because of its flexibility in predicting data trends with configuration changes, proves to be useful in the preliminary design of rocket reaction control systems.

Ice nucleus activity measurements of solid rocket motor exhaust particles

Abstract: The ice Nucleus activity of exhaust particles generated from combustion of Space Shuttle propellant in small rocket motors has been measured. The activity at -20 C was substantially lower than that of aerosols generated by unpressurized combustion of propellant samples in previous studies. The activity decays rapidly with time and is decreased further in the presence of moist air. These tests corroborate the low effectivity ice nucleus measurement results obtained in the exhaust ground cloud of the Space Shuttle. Such low ice nucleus activity implies that Space Shuttle induced inadvertent weather modification via an ice phase process is extremely unlikely.

Oxidizer heat exchanger component testing

Abstract: The RL10-IIB engine, is capable of multimode thrust operation. The engine operates at two low-thrust levels: tank head idle (THI), approximately 1 to 2 percent of full thrust; and pumped idle, 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient thermal conditioning; PI operation provides vehicle tank prepressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-IIB engine during the low-thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidized heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. This report summarizes the test activity and post-test data analysis for two possible heat exchangers, each of which employs a completely different design philosophy. One design makes use of a low-heat transfer (PHT) approach in combination with a volume to attenuate pressure and flow oscillations. The test data showed that the LHT unit satisfied the oxygen exit quality of 0.95 or greater in both the THI and PI modes while maintaining stability. The HHT unit fulfilled all PI requirements; data for THI satisfactory operation is implied from experimental data that straddle the exact THI operating point.

Computational simulation of liquid rocket injector anomalies

Abstract: A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

Heat pipe cooled rocket engines

Abstract: A concept for a heat-pipe-cooled rocket engine, which involves the use of heat pipes to augment radiation cooling on small, low-flux rocket engines used for attitude control and stationkeeping on satellites, is discussed. Results of the thermal and performance analyses performed on a prototype heat cooled thrust chamber are presented, together with the results of fabrication experiments. A test plan for full-scale hot fire hardware testing is outlined.

Controls, Health Assessment, and Condition Monitoring for Large, Reusable, Liquid Rocket Engines

Abstract: The Space Shuttle Main Engine (SSME) represents the state-of-the-art for application of controls, health assessment, and condition monitoring technology to large, reusable, liquid rocket engines. Advances in the technologies that form engine control system along with experience gained in the development and deployment of the SSME point to opportunities for significant improvement in current and future engines. Identification of current engine shortcomings, of future engine requirements, and how control systems technology may be applied to improve life and performance while reducing life cycle costs provides motive for control systems improvement. This paper will derive research and development needs from a survey of liquid rocket control systems technology. The prospects for advanced liquid rocket control systems will be illuminated through discussion of current NASA programs and efforts. Applicable literature is reviewed.