Jeffrey L. Hall

Computational Modeling and Experiments of Natural Convection for a Titan Montgolfiere

Computational models are developed to predict the natural convection heat transfer and buoyancy for a Montgolfiere under conditions relevant to the Titan atmosphere. Idealized single and double-walled balloon geometries are simulated using algorithms suitable for both laminar and (averaged) turbulent convection. Steady-state performance results are compared to existing heat transfer coefficient correlations. The laminar results, in particular, are used to test the validity of the correlations in the absence of uncertainties associated with turbulence modeling. Some discrepancies are observed, especially for convection in the gap, and appear to be primarily associated with temperature nonuniformity on the balloon surface. The predicted buoyancy for the single-walled balloon in the turbulent convection regime, predicted with a standard k ǫ turbulence model, was within 10% of predictions based on the empirical correlations. There was also good agreement with recently conducted experiments in a cryogenic facility designed to simulate the Titan atmosphere.

Simulation and Cryogenic Experiments of NaturalConvection for the Titan Montgolfiere

Natural convection in a spherical geometry is considered for prediction of the buoyancy of single- and double-walled balloons in a cryogenic environment such as Titan’s atmosphere. The steady-state flow characteristics obtained by solving the Reynolds-averaged Navier–Stokes equations with a standard turbulence model are used to determine the net buoyancy as a function of heat input. Thermal radiation effects are shown to have a minor impact on the buoyancy, as would be expected at cryogenic conditions. The predicted buoyancy and temperature fields compare favorably with experiments preformed on a 1-m-diameter Montgolfiere prototype in a cryogenic facility. In addition, both numerical and experimental results were compared with correlations for the heat transfer coefficients for free convection internal and external to the balloon as well as in the concentric gap of the double-walled balloons. Finally, scaling issues related to inferring the performance of the full-scale Montgolfiere from the model-scale results are examined.

FLOW ENERGY PIEZOELECTRIC BIMORPH NOZZLE HARVESTER

There is a need for a long-life power generation scheme that could be used downhole in an oil well to produce 1 Watt average power. There are a variety of existing or proposed energy harvesting schemes that could be used in this environment but each of these has its own limitations. The vibrating piezoelectric structure is in principle capable of operating for very long lifetimes (decades) thereby possibly overcoming a principle limitation of existing technology based on rotating turbo-machinery. In order to determine the feasibility of using piezoelectrics to produce suitable flow energy harvesting, we surveyed experimentally a variety of nozzle configurations that could be used to excite a vibrating piezoelectric structure in such a way as to enable conversion of flow energy into useful amounts of electrical power. These included reed structures, spring mass-structures, drag and lift bluff bodies and a variety of nozzles with varying flow profiles. Although not an exhaustive survey we identified a spline nozzle/piezoelectric bimorph system that experimentally produced up to 3.4 mW per bimorph. This paper will discuss these results and present our initial analyses of the device using dimensional analysis and constitutive electromechanical modeling. The analysis suggests that an order-of-magnitude improvement in power generation from the current design is possible.

Low-Cost Propellant Launch to Earth Orbit from a Tethered Balloon - an Update

As previously reported [1], it may be possible to launch payloads into low-Earth orbit (LEO) at a per- kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch).11 An attractive payload would be large quantities of high-performance rocket propellant as required for the exploration of the moon, Mars, and beyond. The concept is to use small mass-produced rockets that can reach orbit with modest atmospheric drag losses because they are launched from high altitude (e.g. 22 km). These small rockets launch from this altitude by being winched up a tether to a balloon. The drag losses on a rocket are strongly related to the ratio of the rocket launch mass to the mass of the atmospheric column displaced as the vehicle ascends from the launch site to orbit. By reducing the mass of this atmospheric column to a few percent of what it would be launching from sea level, the mass of the rocket can be proportionately reduced while maintaining the drag loss at an acceptably small level.

Low-Cost propellant launch to LEO from a tethered balloon: Recent progress

As we have previously reported [1,2], it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. LO2/LH2) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond. The concept is to use small, mass-produced, two-stage, LO2/LH2, pressure-fed rockets (e.g. without turbo-pumps, which increase performance but are costly). These small rockets can reach orbit with modest atmospheric drag losses because they are launched from very high altitude (e.g. 22 km). They reach this altitude by being winched up a tether to a balloon that is permanently stationed there. The drag losses on a rocket are strongly related to the ratio of the rocket launch mass to the mass of the atmospheric column that is displaced as the vehicle ascends from launch to orbit. By reducing the mass of this atmospheric column to a few percent of what it would be if launched from sea level, the mass of the rocket can be proportionately reduced while maintaining drag loss at an acceptably small level.

Evaluating Low Concept Maturity Mission Elements and Architectures for a Venus Flagship Mission

NASA’s Planetary Science Division recently commissioned a Science and Technology Definition Team to design a potential Venus Flagship mission. The team developed a list of various mission elements that could serve as parts of an overall mission architecture, including orbiters, balloons at various altitudes, and landed platforms of varying number and lifetime. In order to determine the mission architecture that provided the best science within the desired cost range, teams of scientists developed priorities for the science investigations previously detailed by the Venus Exploration Assessment Group (VEXAG). By categorizing the suitability of mission elements to achieve the science investigations, it was possible to construct a Science Figure of Merit (FOM) that could be used to rate the mission elements in terms of their overall science capability. Working in parallel, a team of technologists and engineers identified the technologies needed for the different mission elements, as well as their technology readiness. A Technology FOM was then created reflecting the criticality of a specific technology as well as its technology readiness level. When the Science and Technology FOMs were combined with a rapid costing approach previous developed, it became possible to rapidly evaluate not only individual mission elements, but also their combinations into various mission architectures, accelerating the convergence on a flagship mission architecture that provided the best science within the flagship mission budget, as well as reducing reliance on unproven technology..

Low-cost propellant launch to LEO from a tethered balloon - economic and thermal analysis

As we have previously reported [1–3], it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). 1 2 An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. LO2/LH2) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond. The concept is to use small, mass-produced, two-stage, LO2/LH2, pressure-fed rockets (e.g. without turbopumps, which increase performance but are costly). These small rockets could reach orbit with modest atmospheric drag losses because they are launched from very high altitude (e.g. 22 km). They reach this altitude by being winched up a tether to a balloon that is permanently stationed there. The drag losses on a rocket are strongly related to the ratio of the rocket launch mass to the mass of the atmospheric column that is displaced as the vehicle ascends from launch to orbit. By reducing the mass of this atmospheric column to a few percent of what it would be if launched from sea level, the mass of the rocket could be proportionately reduced while maintaining drag loss at an acceptably small level. The system concept is that one or more small rockets would be launched to rendezvous on every orbit of a propellant depot in LEO. There is only one orbital plane where a depot would pass over the launch site on every orbit - the equator. Fortunately, the U.S. has two small islands virtually on the equator in the mid-Pacific (Baker and Jarvis Islands). Launching one on every orbit, approximately 5,500 rockets would be launched every year, which is a manufacturing rate that allows significantly reduced manufacturing costs, especially when combined with multiyear production contracts, giving a projected propellant cost in LEO of

Low-cost propellant launch to LEO from a tethered balloon — ‘Propulsion depots’ not ‘propellant depots’

400/kg or less. The configuration of the proposed propellant depot and the manner in which the propellant would be utilized has already been reported [1]. The launch processing facility (a small, modified container ship) and cable-car that moves the rocket on the tether have also been reported [2]. This paper provides new analysis of the economics of low-cost propellant launch coupled with dry hardware re-use, and of the thermal control of the liquid hydrogen once on-orbit. One conclusion is that this approach enables an overall reduction in the cost-per-mission by as much as a factor of five as compared to current approaches for human exploration of the moon, Mars, and near-Earth asteroids.

Breakthrough in Mars balloon technology

As we have previously reported [1–4], it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems. The capital investment required would be relatively small, comparable to a single large present-day launch. 1 2 An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond. The concept is to use small, mass-produced, two-stage, LO2/LH2, pressure-fed rockets (without pumps or other complex mechanisms). These small rockets can reach orbit with modest atmospheric drag losses because they are launched from very high altitude (e.g., 22 km). They would reach this altitude by being winched up a tether to a balloon that would be permanently stationed there. The drag losses on a rocket are strongly related to the ratio of the rocket launch mass to the mass of the atmospheric column that is displaced as the vehicle ascends from launch to orbit. By reducing the mass of this atmospheric column to a few percent of what it would be if launched from sea level, the mass of the rocket could be proportionately reduced while maintaining drag loss at an acceptably small level. The system concept is that one or more small rockets would be launched to rendezvous on every orbit of a propellant depot in LEO. There is only one orbital plane where a depot would pass over the launch site on every orbit - the equator. Fortunately, the U.S. has two small islands virtually on the equator in the mid-Pacific (Baker and Jarvis Islands). Launching one on every orbit, approximately 5,500 rockets would be launched every year, which is a manufacturing rate that would allow significantly reduced manufacturing costs, especially when combined with multi year production contracts, giving a projected propellant cost in LEO of

Experimental results for Titan aerobot thermo-mechanical subsystem development

400/kg or less. This paper provides new analysis and discussion of a configuration for the payload modules to eliminate the need for propellant transfer on-orbit. Instead of being a “propellant depot”, they constitute a “propulsion depot”, where propulsion modules would be available, to be discarded after use. The key observation here is that the only way cryo-propellant can get to orbit is by already being in a tank with a rocket engine, and that careful system engineering could ensure that that same tank and engine would be useful to provide the needed rocket impulse for the final application. Long “arms” of these propulsion modules, docked side-by-side, could boost large payloads out of LEO for relatively low-cost human exploration of the solar system.