Albert J. Juhasz

Realizing "2001: A Space Odyssey": Piloted Spherical Torus Nuclear Fusion Propulsion

A conceptual vehicle design enabling fast, piloted outer solar system travel was created predicated on a small aspect ratio spherical torus nuclear fusion reactor. The initial requirements were satisfied by the vehicle concept, which could deliver a 172 mt crew payload from Earth to Jupiter rendezvous in 118 days, with an initial mass in low Earth orbit of 1,690 mt. Engineering conceptual design, analysis, and assessment was performed on all major systems including artificial gravity payload, central truss, nuclear fusion reactor, power conversion, magnetic nozzle, fast wave plasma heating, tankage, fuel pellet injector, startup/re-start fission reactor and battery bank, refrigeration, reaction control, communications, mission design, and space operations. Detailed fusion reactor design included analysis of plasma characteristics, power balance/utilization, first wall, toroidal field coils, heat transfer, and neutron/x-ray radiation. Technical comparisons are made between the vehicle concept and the interplanetary spacecraft depicted in the motion picture 2001: A Space Odyssey.

Design Considerations for Lightweight Space Radiators Based on Fabrication and Test Experience With a Carbon-Carbon Composite Prototype Heat Pipe

Summary This report discusses the design implications for spacecraftradiators made possible by the successful fabrication andproof-of-concept testing of a graphite-fiber-carbon-matrix com-posite (i.e., carbon-carbon (C-C)) heat pipe. The prototype heatpipe, or space radiator element, consists of a C-C compositeshell with integrally woven fins. It has a thin-walled furnace-brazed metallic (Nb- 1%Zr) liner with end caps for containmentof the potassium working fluid. A short extension of this liner,at increased wall thickness beyond the C-C shell, forms theheat pipe evaporator section which is in thermal contact withthe radiator fluid that needs to be cooled. During the fabricationprocess the C_ shell condenser section was exposed to anatomic oxygen (AO) ion source for a total AO fluence of4×102o atoms/cm 2, thereby raising its surface emissivity forheat radiation to a value of 0.85 to 0.90 at design operatingtemperatures of 700 to 800 K. The prototype heat pipe wasextensively tested from startup at ambient conditions, with theworking fluid initially in the frozen state, to a condensertemperature of nearly 700 K. Post-test inspection showed theheat pipe to be in excellent condition after several thermalcycles from ambient to operating temperature.The report also discusses the advantage of segmented spaceradiator designs utilizing heat pipe elements, or segments, intheir survivability to micrometeoroid damage. This survivabil-ity is further raised by the use of condenser sections withattached fins, which also improve the radiation heat transferrate. Since the problem of heat radiation from a fin does not lenditself to a closed analytical solution, a derivation of the govern-ing differential equation and boundary conditions is given inappendix A, along with solutions for rectangular and parabolicfin profile geometries obtained by use of a finite differencecomputer code written by the author.From geometric and thermal transport properties of the C-Ccomposite heat pipe tested, a specific radiator mass of 1.45 kg/m 2 can be derived. This is less than one-fourth the specific massof present day satellite radiators. Using composites with ultra-high conductivity would further reduce the area density ofspacecraft radiators, and utilizing alternate heat pipe fluids withcompatible liner materials would extend the C-C heat pipetechnology to a wide range of temperatures and applications.

A spherical torus nuclear fusion reactor space propulsion vehicle concept for fast interplanetary travel

A conceptual vehicle design enabling fast outer solar system travel was produced predicated on a small aspect ratio spherical torus nuclear fusion reactor. Initial requirements were for a human mission to Saturn with a>5% payload mass fraction and a one way trip time of less than one year. Analysis revealed that the vehicle could deliver a 108 mt crew habitat payload to Saturn rendezvous in 235 days, with an initial mass in low Earth orbit of 2,941 mt. Engineering conceptual design, analysis, and assessment was performed on all major systems including payload, central truss, nuclear reactor (including diverter and fuel injector), power conversion (including turbine, compressor, alternator, radiator, recuperator, and conditioning), magnetic nozzle, neutral beam injector, tankage, start/re-start reactor and battery, refrigeration, communications, reaction control, and in-space operations. Detailed assessment was done on reactor operations, including plasma characteristics, power balance, and component design.A conceptual vehicle design enabling fast outer solar system travel was produced predicated on a small aspect ratio spherical torus nuclear fusion reactor. Initial requirements were for a human mission to Saturn with a>5% payload mass fraction and a one way trip time of less than one year. Analysis revealed that the vehicle could deliver a 108 mt crew habitat payload to Saturn rendezvous in 235 days, with an initial mass in low Earth orbit of 2,941 mt. Engineering conceptual design, analysis, and assessment was performed on all major systems including payload, central truss, nuclear reactor (including diverter and fuel injector), power conversion (including turbine, compressor, alternator, radiator, recuperator, and conditioning), magnetic nozzle, neutral beam injector, tankage, start/re-start reactor and battery, refrigeration, communications, reaction control, and in-space operations. Detailed assessment was done on reactor operations, including plasma characteristics, power balance, and component design.

Development of Lightweight Radiators for Lunar Based Power Systems

This report discusses application of a new lightweight carbon-carbon (C-C) space radiator technology developed under the NASA Civil-Space Technology Initiative (CSTI) High Capacity Power Program to a 20 kWe lunar based power system. This system comprises a nuclear (SP-100 derivative) heat source, a Closed Brayton Cycle (CBC) power conversion unit with heat rejection by means of a plane radiator. The new radiator concept is based on a C-C composite heat pipe with integrally woven fins and a thin walled metallic liner for containment of the working fluid. Using measured areal specific mass values (1.5 kg/m2) for flat plate radiators, comparative CBC power system mass and performance calculations show significant advantages if conventional heat pipes for space radiators are replaced by the new C-C heat pipe technology.

HIGH CONDUCTIVITY CARBON-CARBON HEAT PIPES FOR LIGHT WEIGHT SPACE POWER SYSTEM RADIATORS

Abstract Based on prior successful fabrication and demonstration testing of a carbon-carbon heat pipe radiator element with integral fins this paper examines the hypothetical extension of the technology via substitution of high thermal conductivity composites which would permit increasing fin length while still maintaining high fin effectiveness. As a result the specific radiator mass could approach an ultimate asymptotic minimum value near 1.0 kg/m 2 , which is less than one fourth the value of present day satellite radiators. The implied mass savings would be even greater for high capacity space and planetary surface power systems, which may require radiator areas ranging from hundreds to thousands of square meters, depending on system power level. Nomenclature A i inner surface area; A o outer surface area A r ( j ) incremental radiator area at section j dA i elemental inner wall surface, or heat pipe evaporator area dA r dA o , is the elemental radiating outer wall surface, or heat pipe condenser area

Testing of a Liquid Metal Carbon‐Carbon Heat Pipe

Proof‐of‐concept testing of a potassium‐filled, carbon‐carbon (C–C) heat pipe with integrally woven fins has been completed. The objectives of these tests were to verify mechanical stability of the pipe through multiple thermal cycles and to assess heat pipe performance. The heat pipe was operated for 11 h, starting at ambient temperature with the potassium working fluid in a frozen state. Smooth transition to steady‐state conditions was achieved; however, maximum power was limited by the electrical heater capability. Heat pipe output power, as measured by a water calorimeter, was 300 W at 700 K, which represents approximately 50% of the maximum design power. Additional testing at higher power levels and temperatures are planned. Posttest inspection of the heat pipe showed the C‐C structure to be in excellent condition after eight thermal cycles from ambient temperature. No evidence of liner debonding or deformation was observed and containment of potassium within the niobium‐1% zirconium (Nb‐lZr) boundar...

An analysis and procedure for determining space environmental sink temperatures with selected computational results

The purpose of this paper is to analyze the heat transfer problem posed by the determination of spacecraft temperatures and to incorporate the theoretically derived relationships into a computational code. Subject code is based on a theoretical analysis of thermal radiative equilibrium in space, particularly in the Solar System. Starting with the solar luminosity, the code takes into account a number of key variables, namely; the spacecraft-to-Sun distance expressed in AU (Astronomical Units), with, 1 AU representing the average Sun-to-Earth distance of 149.6 million km; the angle (degrees of arc) at which solar radiation is incident on a spacecraft surface, the temperature of which is to be determined (i.e., a radiator or PV (photovoltaic) array); the absorptivity-to-emissivity ratio of the surface, /spl alpha///spl epsiv/, with respect to solar radiation; and the view factor of the surface to space.

Carbon-carbon heat pipe testing and evaluation

This report discusses development and proof-of-concept testing of a new lightweight carbon-carbon (C-C) space radiator heat pipe developed under the NASA Civil Space Technology Initiative (CSTI) High Capacity Power Program. The heat pipe was filled with potassium working fluid and tested for 11 hours including startup from ambient temperature with the working fluid initially in the frozen state to near 700 K condenser temperature. Steady-state heat pipe input power during testing was facility limited to about 300 watts, representing about 50 percent of the design input power. Post test inspection showed the heat pipe to be in excellent condition after eight thermal cycles from ambient to steady-state operating temperature. Potential applications, ranging from small spacecraft heat rejection to aircraft and terrestrial uses, are discussed.

Development of lightweight prototype carbon‐carbon heat pipe with integral fins and metal foil liner

This report discusses development and proof‐of‐concept testing of a new lightweight carbon‐carbon (C‐C) space radiator heat pipe developed under the NASA Civil Spqce Technology Initiative (CSTI) High Capacity Power Program. The heat pipe was filled with potassium working fluid and tested for 11 hours, including startup from ambient temperature with the working fluid initially in the frozen state to near 700 K condenser temperature. Steady‐state heat pipe input power during testing was facility limited to about 300 watts, representing about 50% of the design input power. Post test inspection showed the heat pipe to be in excellent condition after eight thermal cycles from ambient to steady‐state operating temperature. Potential applications, ranging from small spacecraft heat rejection to aircraft and terrestrial uses, are discussed.

Mathematical Analysis of Space Radiator Segmenting for Increased Reliability and Reduced Mass

Spacecraft for long duration deep space missions will need to be designed to survive micrometeoroid bombardment of their surfaces some of which may actually be punctured. To avoid loss of the entire mission the damage due to such punctures must be limited to small, localized areas. This is especially true for power system radiators, which necessarily feature large surface areas to reject heat at relatively low temperature to the space environment by thermal radiation. It may be intuitively obvious that if a space radiator is composed of a large number of independently operating segments, such as heat pipes, a random micrometeoroid puncture will result only in the loss of the punctured segment, and not the entire radiator. Due to the redundancy achieved by independently operating segments, the wall thickness and consequently the weight of such segments can be drastically reduced. Probability theory is used to estimate the magnitude of such weight reductions as the number of segments is increased. An analysis of relevant parameter values required for minimum mass segmented radiators is also included.