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Brief report proposes use of jet boost pumps in conjunction with main pumps supplying liquid hydrogen and liquid oxygen to main engine of Space Shuttle. Main part of pump has no moving parts. Benefits include increased reliability, simplified ducts, and decreased weight.
Gun barrel erosion is serious problem with two-stage light gas guns. Excessive barrel erosion can lead to poor or failed launches and frequent barrel changes, with the corresponding down time. Also, excessive barrel erosion can limit the maximum velocity obtainable by loading down the hydrogen working gas with eroded barrel material. Guided by a CFD code, the operating conditions of the Ames 0.5-inch gun were modified to reduce barrel erosion. The changes implemented included: (1) reduction in the piston mass, powder mass and hydrogen fill pressure; and (2) reduction in pump tube volume, while maintaining hydrogen mass. The latter change was found, in particular, to greatly reduce barrel erosion. For muzzle velocity ranges of 6.1 - 6.9 km/sec, the barrel erosion was reduced by a factor of 10. Even for the higher muzzle velocity range of 7.0 - 8.2 km/sec, the barrel erosion was reduced by a factor of 4. Gun erosion data from the Ames 0.5-inch, 1.0-inch, and 1.5-inch guns operated over a wide variety of launch conditions was examined and it was found that this data could be correlated using four different parameters: normalized powder charge energy, normalized hydrogen energy density, normalized pump tube volume and barrel diameter. The development of the correlation and the steps used to collapse the experimental data are presented. Over a certain parameter range in the correlation developed, the barrel erosion per shot is found to increase very rapidly. The correlation should prove useful in the selection of gun operating conditions and the design of new guns. Representative shapes of eroded gun barrels are also presented.
Cold air performance of scale model oxidizer pump drive turbine feedpipe-manifold-nozzle assembly for M-1 rocket engine
Atomic-based-combined-cycle (ABCC) engine combines an air-breathing ramjet engine with an atomic reactor to increase the mission-averaged specific impulse and thereby increasing the dry-mass ratio. ABCC engine is similar to RBCC engine except that energy needed for the propulsive power is derived from nuclear reaction rather than chemical combustion used in the RBCC engine. The potential performance improvement of an ABCC engine over a RBCC engine comes from two factors. Firstly, the energy density of nuclear reaction is several order of magnitudes higher than the chemical combustion. Secondly, hydrogen can produce much higher nozzle exit velocity because of its small molecular weight. A one-dimensional, transient numerical model was used to analyze a generic scramjet engine and it is used as a baseline to evaluate an imaginary ABCC engine performance. A nuclear reactor is treated as a black box energy source that replaces the role of the primary rocket and the chemical combustion chamber in a RBCC engine. Hydrogen is heated by the reactor and accelerated to produce high-speed ejection velocity. The ejection velocity up 10,000 m/sec is theoretically possible because of high energy density from the reactor and large gas constant of the hydrogen. Oxygen contained in the entrained air reacts with hydrogen and produces propulsive power for ejector mode operation. To provide enough thrust for initial acceleration, relatively large amount of hydrogen must be pumped through the reactor. Amount of oxygen contained in the entrained air may not be sufficient to burn all hydrogen and consequently combustion could occur at the end of exit nozzle. It is assumed that this combustion process is constant-pressure combustion at 1.0 atmospheric pressure and thus not affects the nozzle exit condition.
Vacuum pumps are the heart of a fusion energy facility – fusion power cannot be generated without them. Nevertheless, the vacuum technology needed to operate a viable fuel cycle for a compact fusion power plant does not exist. Commercial vacuum technology offers the best solution to this challenge, but a pump oil recycling and detritiation system is necessary. Conventional oil detritiation processes have only been developed to deal with legacy waste and are too slow and destructive to the oil. Further, post hoc detritiation strategies are intrinsically inefficient. Our approach is to rethink the challenge holistically by designing the pumping fluid and tritium extraction system in an integrated manner to achieve an innovative solution. By selecting an oil for the specific task of pumping tritium and then designing our catalytic system to selectively target the tritium-bearing functional groups, an effective detritiation system (hydrocarbon pump oil recycling, HyPOR, loop) for fusion power plants can be obtained. The project demonstrated a HyPOR loop process that can selectively remove heavier hydrogen isotopes from pump oil, reaching the target of 99.5 % removal, with an uptake of less than 0.01% of tritium throughput, while also purifying the oil of radiation-induced damage. The recycled oil retained its pumping characteristics over 7 recycles and gamma irradiation over 30 MGy. By meeting these targets, the project enable a reduction in pump operational costs from $\$$14.5M/year to $\$$103k/year (>140× reduction), reducing pump electric power consumption from 2.8 MW to 0.25 MW (>10× reduction), and reducing in-process tritium inventory from 2.03 kg to 0.48 kg (>4× reduction).
Of the six chilldown tests, data from only one could be used for evaluation. During the rest of the chilldown tests, there was leakage hydrogen flow into the pump cavity prior to the initiation of the chilldown test. In all of the tests the hydrogen condition into the pump was probably 100% vapor. The data from this one test, therefore, can be used to compare only the single phase fluid correlation in the analytical pump chilldown model. In general, the actual pump chilled down much faster than predicted by the analytical pump model. There were insufficient data from the test to measure the pump flow rate and pump inlet fluid condition; therefore, these parameters were extrapolated based on related data which were available. However, even with the highest probable flow rate, the pump chilled faster than predicted.
Cold air performance test of scale model oxidizer pump drive turbine for M-1 engine - performance of first stage with inlet feedpipe manifold assembly
Aerodynamic performance of scale model oxygen pump-drive turbine for M-1 rocket engine
Studies to define the optimum turbopump configurations for the Space Shuttle propulsion system are reported. For each propellant, two turbopumps - one low-pressure and one high-pressure - are needed to generate the high discharge pressures required for engine operation. The optimization of the four pumps resulted in the selection of an axial inducer type for both low-pressure pumps and in a three-stage centrifugal pump for the high-pressure hydrogen turbopump; meanwhile the oxygen pump features a double inlet main impeller. This impeller feeds the thrust chamber injector, as well as a preburner pump, which boosts the main impeller discharge pressure to the required preburner pressure. The interaction between engine performance and obtainable turbopump performance is discussed, and the reasons for the final selection of the turbine and pump types are presented.
Pump Design (PUMPDES) is a computer program for designing a rotating pump for liquid hydrogen, liquid oxygen, liquid nitrogen, water, methane, or ethane. Using realistic properties of these fluids provided by another program called GASPAK, this code performs a station-by-station, mean-line analysis along the pump flow path, obtaining thermodynamic properties of the pumped fluid at each station and evaluating hydraulic losses along the flow path. The variables at each station are obtained under constraints that are consistent with the underlying physical principles. The code evaluates the performance of each stage and the overall pump. In addition, by judiciously choosing the givens and the unknowns, the code can perform a geometric inverse design function: that is, it can compute a pump geometry that yields a closest approximation of given design point. The code contains two major parts: one for an axial-rotor/inducer and one for a multistage centrifugal pump. The inducer and the centrifugal pump are functionally integrated. The code can be used in designing and/or evaluating the inducer/centrifugal-pump combination or the centrifugal pump alone. The code is written in standard Fortran 77.
This paper describes recent developments in gun operating techniques at the Ames ballistic range complex. This range complex has been in operation since the early 1960s. Behavior of sabots during separation and projectile-target impact phenomena have long been observed by means of short-duration flash X-rays: new versions allow operation in the lower-energy ("soft") X-ray range and have been found to be more effective than the earlier designs. The dynamics of sabot separation is investigated in some depth from X-ray photographs of sabots launched in the Ames 1.0 in and 1.5 in guns; the sabot separation dynamics appears to be in reasonably good agreement with standard aerodynamic theory. Certain sabot packages appear to suffer no erosion or plastic deformation on traversing the gun barrel, contrary to what would be expected. Gun erosion data from the Ames 0.5 in, 1.0 in, and 1.5 in guns is examined in detail and can be correlated with a particular non- dimensionalized powder mass parameter. The gun erosion increases very rapidly as this parameter is increased. Representative shapes of eroded gun barrels are given. Guided by a computational fluid dynamics (CFD) code, the operating conditions of the Ames 0.5 in and 1.5 in guns were modified. These changes involved: (1) reduction in the piston mass, powder mass and hydrogen fill pressure and (2) reduction in pump tube volume, while maintaining hydrogen mass. These changes resulted in muzzle velocity increases of 0.5-0.8 km/sec, achieved simultaneously with 30-50 percent reductions in gun erosion.
A liquid lithium divertor can improve performance of future fusion devices by creating efficient power exhaust and improving the energy confinement via pumping of the hydrogen isotopes. In addition, significantly higher heat fluxes can be handled if controlled vapor shielding is used to redistribute the divertor heat flux over a wider area. Design and optimization of such a system calls for an analysis model which includes a strong two-way coupling between the plasma and divertor material. The incoming plasma heat and particle flux will affect the divertor surface temperature, which is a defining factor of the lithium evaporative and sputtered flux going into the plasma. Results of the coupled model based on the plasma edge code SOLPS-ITER and the computational fluid dynamics (CFD) code ANSYS-CFX will be presented for different configurations. An analytical slab flow model is used as a heat transfer boundary condition for SOLPS, defining particle flux from the wall via calculation of the surface temperature. At the final step, results of the SOLPS analysis are verified using a 3D CFD magnetohydrodynamics (MHD) analysis which uses heat and particle flux from SOLPS as a boundary condition. In addition to plasma heat flux, both analytical and CFD temperature models include several plasma material interaction effects, such as lithium evaporation, condensation and sputtering based on deuterium target flux. New adatom sputtering model based on the available experimental data is presented. Analytical model is expanded to include free surface axisymmetric configurations. Results of parametric studies of the divertor configurations with different lithium inlet temperature and velocity will be presented leading to the optimal design resulting in the lowest possible lithium contamination in the core.
In support of the U.S. Department of Energy’s Energy Storage Grand Challenge, Pacific Northwest National Laboratory is applying its rich history of battery research and development to provide the U.S. Department of Energy (DOE) and industry with a guide to current energy storage costs and performance metrics for various technologies. Cost and performance metrics for individual technologies track the following to provide an overall cost of ownership for each technology: cost to procure, install, and connect an energy storage system; associated operational and maintenance costs; and end-of-life costs. These metrics are intended to support DOE and industry stakeholders in making sound decisions about future R&D directions and priorities that move the United States closer to its goal of energy independence. The technologies currently being evaluated are lithium-ion [lithium iron phosphate (LFP) and nickel manganese cobalt (NMC)] batteries, vanadium redox flow batteries, lead-acid batteries, zinc-based batteries, hydrogen energy storage, pumped storage hydropower, gravitational energy storage, compressed air energy storage, and thermal energy storage.
Requirements for the orbit transfer vehicle engine were examined. Engine performance/weight sensitivities, the effect of a service life of 300 start/shutdown cycles between overalls on the maximum engine operating pressure, and the sensitivity of the engine design point (i.e., thrust chamber pressure and nozzle area ratio) to the performance requirements specified are among the factors studied. Preliminary engine systems analyses were conducted on the stage combustion, expander, and gas generator engine cycles. Hydrogen and oxygen pump discharge pressure requirements are shown for various engine cycles. Performance of the engine cycles is compared.
The axial pump is considered in terms of the total turbopump assembly. Stage hydrodynamic design, pump rotor assembly, pump materials for liquid hydrogen applications, and safety factors as utilized in state of the art pumps are among the topics discussed. Axial pump applications are included.
Two hydrogen oxygen rocket engine system concepts were analyzed parametrically over a thrust range from 100 to 1000 pounds and a chamber pressure range from 175 to 1000 psia. Both concepts were regeneratively cooled with hydrogen and were pump fed by electric motor driven positive displacement pumps. Electric power was provided by either a turboalternator (turboalternator concept) or some means external to the engine system (auxiliary power concept). The turboalternator concept is discussed. The computer program used to conduct the analyses along with the design characteristics of the major engine system components is described. The feasible design range of the systems over the parametric range of thrust is discussed in terms of allowable chamber pressure. Engine system estimated performance, mass, and dimensional envelope parametric data within the feasible design range are presented.
A study was conducted within the thrust range 450 to 9000 N (100 to 2000 pounds). Performance analyses were made on centrifugal, pitot, Barske, drag, Tesla, gear, piston, lobe, and vane pumps with liquid hydrogen, liquid methane, and liquid oxygen as propellants. Gaseous methane and hydrogen driven axial impulse turbines, vane expanders, piston expanders, and electric motors were studied as drivers. Data are presented on performance, sizes, weights, and estimated service lives and costs.
This paper presents the results of a combined analytical and experimental investigation of low specific speed pumps for potential use as components of propellant feed systems for low thrust rocket engines. Shrouded impellers and open face impellers were tested in volute type and vaned diffuser type pumps. Full- and partial-emission diffusers and full- and partial-admission impellers were tested. Axial and radial loads, head and efficiency versus flow, and cavitation tests were conducted. Predicted performance of two pumps are compared when pumping water and liquid hydrogen. Detailed pressure loss and parasitic power values are presented for two pump configurations. Partial-emission diffusers were found to permit use of larger impeller and diffuser passages with a minimal performance penalty. Normal manufacturing tolerances were found to result in substantial power requirement variation with only a small pressure rise change. Impeller wear ring leakage was found to reduce pump pressure rise to an increasing degree as the pump flowrate was decreased.