US20080164012A1

US20080164012A1 - Radiative cooling system

Info

Publication number: US20080164012A1
Application number: US11/796,528
Authority: US
Inventors: David B. Cavanaugh; Richard S. Dummer; Steven L. Palm
Original assignee: Surface Optics Corp
Current assignee: Surface Optics Corp
Priority date: 2006-12-11
Filing date: 2007-04-26
Publication date: 2008-07-10

Abstract

A radiator unit for radiating heat from a heat source into space. The radiating unit includes an inflatable radiating panel comprising a plurality of inflatable tubes inflated with and adapted to contain and permit a circulation through the plurality of inflatable tubes of a low mass gas, a sufficient quantity of low mass gas to inflate said plurality of inflatable tubes and a blower unit for forcing the circulation of said low mass gas through the heat source and said plurality of inflatable tubes.

Description

This invention relates to electric cooling systems and in particular to radiative cooling systems for applications in space. This application is a continuation-in-part of Ser. No. 11/636,919, filed Dec. 11, 2006 which is incorporated by reference herein.

BACKGROUND OF THE INVENTION

Heat is removed from hot things by conduction, convection or radiation. In outer space equipment is typically isolated so the places to which heat can be conducted is very limited. There is no air so convection is out of the question. Basically, any heat generated in equipment located outer space must be radiated away. In general accumulated heat results in an increase in the temperature around whatever is producing the heat until the heat removed equals the heat produced and at that temperature equilibrium is established. The solution in outer space is to establish systems so that the temperature at this equilibrium temperature in not too high. This is typically not a problem when the rate of heat production is very low. However, when very large amounts of heat are produced special equipment must be provided to radiate the heat away.
There are a large number of satellites, both military and commercial, currently in orbit around the earth. Other space vehicles are in use exploring the solar system. In the future many additional space vehicles will be built and used to provide communication and surveillance and other needs here on the earth and for exploring the universe. All of these systems need electric power and many will require special system to get rid of unwanted heat. Unwanted heat is produced when the electric power is generated and when it is utilized.
With respect to electric power generation, most space vehicles currently use solar voltaic systems to convert sunlight into electric power. Heat sinks are generally not required for solar voltaic systems. As the power requirements approach 100 kilowatts, the size of a solar voltaic array becomes impractical. Other systems for generating electric power such as nuclear reactors require a large heat sink. Thermoelectric power generators utilizing radioactive heat sources may also require a special heat sink.
What is needed is better way to safely radiate away heat in outer space.

SUMMARY OF THE INVENTION

The present invention provides radiator unit for radiating heat from a heat source into space. The radiating unit includes an inflatable radiating panel comprising a plurality of inflatable tubes inflated with and adapted to contain and permit a circulation through the plurality of inflatable tubes of a low mass gas, a sufficient quantity of low mass gas to inflate said plurality of inflatable tubes and a blower unit for forcing the circulation of said low mass gas through the heat source and said plurality of inflatable tubes.
The inflatable panel can be compacted into a small volume package for transport into space. The low mass gas is transported into space in one or more gas bottles at very high pressure. In preferred embodiments the low-mass gas is hydrogen and the radiating panel includes tubes with a thin metal lining to make the tubes virtually impermeable to hydrogen.
The radiator unit has numerous applications. In a preferred embodiment the unit functions as a secondary heat transfer loop in which heat from a steam turbine exhaust is transferred from steam to the hydrogen gas through a conventional cross-flow heat exchanger. The hydrogen is pumped through the radiator tubular structure where the heat is radiated to space from the surfaces of the radiating panel. This heat transfer process condenses the steam to water which, with solar heat, is turned to steam to drive the steam turbine. The high thermal heat capacity of the hydrogen makes it an excellent heat transfer medium for this purpose. In this preferred embodiment only 18 kilograms of hydrogen is needed for the radiator system.
This radiator system when deflated can be rolled or folded into a lightweight, compact and rugged bundle. When in space the structure is expanded by inflation with high-pressure hydrogen carried into space in one or more gas bottles. The preferred material for the inflatable structure is a thermoset polyimide or a polyethylene terepthalate, such as Kapton or Mylar, which is lined with an aluminum barrier layer that makes the structure virtually impermeable to hydrogen.
Preferred applications of the present invention also include cooling large and dense bays of electronic components that are mounted on thermally conductive plates, which transfer their heat to a set of heat transfer fins by which hydrogen is circulated. In these applications hydrogen is the preferred working fluid except possibly when the potential presence of oxygen poses an explosion hazard. This is generally not a problem in most space system since the amount of oxygen available is minimal. However, if there is an oxygen risk, the risk can be avoided by utilizing an alternate primary fluid (such water, helium or nitrogen) in regions where oxygen may be present. Hydrogen may then be used as a secondary fluid. This however requires a heat exchanger to transfer the heat from the primary fluid to the hydrogen. Another option is to replace the hydrogen with a non-combustible gas such as nitrogen or helium.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a basic schematic of a preferred embodiment of the present invention.

FIG. 2 shows the spectral reflectance of a preferred balloon coating candidate.

FIG. 3 shows a preferred reflector design.

FIG. 4 shows a communication satellite powered by an embodiment of the present invention.

FIGS. 5A and 5B show features of a heat exchanger system to provide a working gas for radiator units.

FIGS. 6A and 6B show features of the radiator units.

FIGS. 7A and 7B are drawings showing features of a preferred radiating panel.

FIG. 8 shows pumping power as a function of hydrogen flow.

FIG. 9 shows radiating area as a function of gas flow and pressure.

FIG. 10 shows radiator mass as a function of exit temperature.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

First Preferred Embodiment

Radiator for Electric Steam Generator Powering a Satellite

In a preferred embodiment, the present invention is utilized as a heat radiating unit providing a heat sink for a steam driven turbine generator providing electric power for an earth satellite. The entire electric generating system is described in the parent patent application referred to above and much of that described is repeated below. Heat from a steam turbine exhaust is transferred from steam to hydrogen gas through a conventional cross-flow heat exchanger. The hydrogen is pumped through a tubular radiating panel where the heat is radiated to space from the surfaces of the radiating panel. This heat transfer process condenses the steam to water which, with solar heat, is turned to steam to drive the steam turbine. The high thermal heat capacity of the hydrogen makes it an excellent heat transfer medium for this purpose. In this preferred embodiment only 18 kilograms is needed for the radiator system. The hydrogen gas absorbs the heat from the primary working fluid in a cross-flow heat exchanger as shown in FIGS. 5 and 7. The heated gas is pumped though a “radiator panel” 30 consisting of an array of parallel tubes 31 that joined inlet manifold 32A and outlet manifold 32B as shown in FIGS. 6A and B. The heat contained by the gas is transferred to the walls of the tubing as the gas transits their length, and from there is radiated to space.
The total area required of the radiator panels is a function of the thermodynamic properties of the working gas and the flow rate and pressure in the radiator panels. See FIGS. 8, 9 and 10. Applicants' current estimates are that a radiator area of 700 square meters will be required for each of two systems. The material currently envisioned for the radiator panel array is Kapton, coated with a thin aluminum layer to minimize gas permeation.
A schematic diagram of the radiator system is shown in FIG. 7. The performance of the radiator system was modeled using a finite volume code for modeling flow, heat transfer and thermal radiation. Conservation of energy in the simulation was performed in the integral sense. The simulation models the full compressible gas dynamics. The simulation did not assume an ideal or thermally perfect gas, but uses real gas data interpolated from the NIST Web-book thermodynamic tables. The channel exit pressure is left as a free variable simplifying the solution by allowing marching versus using an implicit solver. Local compressible gas dynamics and heat transfer are solved using iterative solvers with under-relaxation. Friction factors are derived using the classical laminar relation for Re<2300 and the PKN (Prandtl (1944), von Karman (1934), Nikuradse (1932)) relation for Re>2300. A smooth transition is made between the two relations in the region from 2100<Re<2500. This transition is smooth to 2nd order. Local Nusselt numbers are obtained using relations by Gnielinski (1976). Radiation view-factors between neighboring components are considered. Some selected results of the simulation and a baseline configuration for the radiator for the power system are presented in FIGS. 8, 9 10 and Table 1 below:


Selected Conditions for GHB Power System:

	Channel inlet Mach number:	0.004
	Channel exit temperature:	260 K
	Operating pressure:	7 atm

Radiator Results:

	Mass Flow Rate:	0.84 kg/sec
	Channel Diameter:	1.8 cm
	Number Channels:	1119
	Channel Length:	46.4 m
	Channel Velocity:	5.4 m/sec
	Channel ReD:	6200
	Channel NuD:	19.6
	Total Frontal Area:	981 m2
	Compressor Power Requirement:	1.7 kW
	Total Radiator Mass:	578 kg

	(Does not include compressor or HX)

Satellite Turbine Generator Power System

As stated above the present invention may be used as the heat sink component in a solar heated steam turbine generator system for a satellite. Such a steam turbine generator system is referred to by the Applicant as “The Greenhouse Balloon Power System” (referred to in this detailed description as “GHB”. It uses an inflatable structure to collect and transfer solar energy to a pressurized working fluid inside the structure. The pressurized working fluid drives a turbo generator producing electrical power. The heated vapor that has traversed the turbine is condensed in a heat exchange element, and the waste heat is radiated to space. The condensed working fluid is recycled into the Greenhouse Balloon. The basic operational scheme for the GHB system is illustrated in FIG. 1. The GHB collectors generate pressurized saturated vapor which spins a turbine/electrical generator combination. The current preferred thermodynamic cycle of the turbine is the Rankine cycle, although other compatible thermodynamic cycles can also be used.

Satellite Power System

The steam turbine generator provides power for a communication satellite is shown in FIG. 4. The satellite consists of two GHB power systems 6 arranged symmetrically around the antenna/transmitter payload 8. The GHB collector panels 10 are shown vertical with the reflector panels 12 directly behind them. The turbine and generator (two sets), with the pressure feed and return lines are located above the payload antenna. The thermal radiator panels are perpendicular to, and extend behind the GHB collector arrays. The pointing directions of the collector panels and the antenna are controlled separately.
The current system uses two symmetrical GHB power generation systems. The reason for this is the solar collector arrays have a large area that points into the solar wind. A system with an unsymmetrical aspect ratio will experience a torque, and a secondary propulsion system will be necessary to correct the attitude of the spacecraft. The symmetrical shape shown here does not generate a torque in the solar wind, and the complexity of the attitude control system is thus minimized.

Solar Collectors

The GHB solar collectors consist of reinforced cylindrical bladders that are designed to hold hot vapor at pressures of the order of 100 psia and temperatures on the order of 160° C. The GHB bladders are flexible when not inflated and can be folded or rolled into small volumes for stowage and launch. Thus, the entire system is designed to fit into the fairing of an Atlas 5 HLV rocket.
The bladders can be arranged in various configurations. The current embodiments arrange the bladders in a parallel structure. The bladders are separated by a minimum distance that prevents the bladders from “shadowing” its neighbors when the GHB raft is kept within a specified angle to the sun. The expected orbital attitude variations determine this specified angle. The bladders are interconnected so that all are connected either in parallel or series to the turbine inlet. The interconnect fittings would include check valves or other safety features at each bladder designed to disconnect the bladder from the system in case of a failure.
In a preferred embodiment shown in FIG. 4, the solar collector is two rectangular panels, each consisting of twenty-five 0.5 meter diameter tubes 20 meters in length spaced at 1.9 meters. These panels therefore are each provide a surface area of about 1152 square meters and an absorption area of about 250 square meters. The two panels together present an absorption area of about 500 square meters. Since the exo-atmospheric solar power is about 1.368 kW/m², this is a sufficient absorption area to receive about 640 kW. With a Rankine cycle steam turbine operating at a projected efficiency of 20 percent and considering collection efficiencies, the system can produce about 100 kW of electrical power. The reader should note that the FIG. 4 drawing shows only 13 tubes per panel instead of the above 25. This may be the number needed if the reflector contribution turns out to be as effective as hoped.
As stated above the outer surface of the GHB tubular collector elements are treated with a coating that selectively absorbs solar radiation with high efficiency. The coating is also designed to retain heat energy in the GHB by having a very low emissivity (i.e. a high reflectivity) in the thermal infrared waveband. Thus, the solar electromagnetic radiation is absorbed by the GHB skin where it is immediately converted to heat energy. The heat energy is trapped very efficiently by the GHB because the outward radiative loss pathway is blocked by the low emissivity coating, and because the outward convective heat loss pathway is not effective in the vacuum of space. The trapped heat thus diffuses to the interior of the GHB where it is absorbed by the working fluid. The coating performance is illustrated in FIG. 2. The coating on the outer surface of the GHB is absorptive to solar light (left side of chart) at about 65 percent and highly reflective to infrared radiation (right side of chart).

Making the Collector Tubes

The GHB collection tubes are preferentially cylindrical in shape to facilitate manufacture. The GHB material is a fiber-reinforced composite as explained above. The fiber reinforcement should have a low mass density, be flexible with a high tensile strength that retains its strength when exposed to steam. Kevlar is the currently the preferred fiber, but other high tensile polymer fibers are currently under study. The Kevlar is applied to the GHB by continuous winding onto a cylindrical mandrel, which forms a seamless material. The fiber is wound both in the “hoop” direction around the cylinder circumference, and longitudinally, parallel with the cylinder length. The fiber is impregnated with the matrix rubber material before winding to eliminate voids in the finished material. The matrix of the composite is a rubbery material that is flexible, withstands high temperature and is impermeable to steam. The current preferred material is a “Viton” blend, although several other rubber candidate materials are currently under study. The thickness of the GHB composite skin is approximately 10 mils (0.01″).
The outer surface of the GHB has a thin film coating that absorbs sunlight and has low emissivity in the thermal infrared. The current preferred coating is composed of three layers that are applied to the surface by vacuum deposition. The base layer is 180 Angstroms if TiO₂. The middle layer is 180 Angstroms of metallic silver. The top layer is another 180 Angstrom layer if TiO₂. This coating has been successfully applied to several samples of the various rubber materials that are under study, and give the performance indicated in FIG. 2. The rubber surface may also be pre-treated by depositing thin layers to the rubber surface, for example a chromium metal, which improves the coating adhesion and the solar absorption properties. Other coatings are also currently being explored, such as black oxides of metals such as chromium or copper.
The interior surface of the GHB is a thin metal foil that blocks permeation of the working fluid through the GHB skin and spreads heat laterally. This metal must be chemically inert to steam, and thus a common metal such as aluminum is unsuitable. The metal foil is also the base layer that the “wet” components of the GHB cylinder are applied to during manufacture. Applicants currently prefer manganese although several options are available including some non-metals and composites.
The preferred diameter of the GHB cylinders is determined by several factors. The effective collection area of the GHB cylinders is equal to the (diameter×length). The GHB elements in the current embodiment must collect solar power to provide the desired electrical power output. The turbine efficiency is determined by the thermodynamic cycle used, and in the case of the Rankine cycle an efficiency of 20 percent might be expected. The exo-atmospheric solar power is 1,368 kW/meter².
The optimal diameter of the GHB is a complex function. A larger diameter GHB gives a large collection area per unit length of GHB cylinder, which is desirable. However, as the diameter of a pressure vessel increases the hoop tensile strength required of the walls increases as the square of the hoop radius. Thus, a smaller diameter GHB will have thinner walls, and thus a lower specific mass.

Reflector

The GHB collector panels preferably are fitted with a reflector screen that is mounted “behind” the GHB panels as shown at 18 in FIG. 3. This screen is made of a lightweight flexible material coated with a reflective layer that is specular to solar wavelengths. The reflector screen is configured with supports and angular ridges as shown at 20 in FIG. 3 so that it reflects the solar energy that misses the GHB collectors (i.e. passes between the GHB collectors) back onto the sides and rear surfaces of the GHB collectors. The reflector also functions as an insulator that helps trap the small level of thermal energy emitted from the dark side of the GHB collector. The reflector also shields the radiator from direct sunlight.
The preferred reflector panels are comprised of a thin sheet of mylar that has a thin layer of aluminum deposited on it by vacuum deposition. Similar ultra-lightweight sheet material is manufactured and deployed for solar sails in spacecraft. A lightweight expandable mechanical structure for spacing and supporting the GHB bladders will also shape and support the reflector sheets. The reflector increases the effective collection area of the GHB, and the properties, design, and construction of the reflector determine the degree of increase. In the ideal reflector the GHB collection area is approximately equal to (GHB area+reflector area), but the anticipated system performance is somewhat below this ideal. The actual efficiencies will be determined through experiment.

Thermal Capacity

The thermal capacity of the GHB is important because during the eclipse periods where the satellite is shadowed from the sun by the earth, thermal energy can still be extracted from the hot GHB system which can be used to maintain the system during the eclipse period. The amount of available energy is a function of the mass of hot working fluid contained in the system, and this mass is limited by the mass budget at satellite launch.
The fluid reservoir mass is adjustable through the GHB diameter in the following manner. Given two GHB solar collection systems with equal collection area, the GHB system with larger diameter cylinders will contain a greater mass of working fluid than a GHB system with smaller diameter cylinders. The optimal GHB diameter is thus determined through a trade study. The GHB wall thickness and thus system mass is reduced by decreasing the GHB diameter, and a sufficient thermal reservoir to carry the system through the eclipse period is insured by increasing the GHB diameter. The outcome of the trade should provide a minimal mass for the GHB collector system, and a working fluid mass that fits within the launch vehicle's mass budget but provides the required thermal reservoir during eclipse.

Working Turbine-Generator Fluid

The primary working fluid used in the turbine generator system is currently pure water, which has several advantages. Water has a relatively low molecular weight which minimizes the mass that must be launched. Water boils in a temperature range in which materials are available that can contain it. Also, water condenses at a high enough temperature that a radiator with a surface area of reasonable size is feasible. Other potential working fluids such as ammonia or freon boil and condense at temperatures that are too low for exhausting excess heat efficiently by radiation. The use of mixtures of working fluid, such as water and ammonia may have potential for improving the system efficiency, such as in the Kalina thermodynamic cycle. Such mixtures of working fluid have not been fully explored, but are under consideration.
The amount of working fluid required to charge the system is determined by the total volume of the system, including the GHB and associated plumbing and heat exchanger. The optimum mass charge will generate the working pressure of saturated vapor in the system volume when raised to the working temperature. The working temperature and pressure are determined by the turbine cycle. In the present embodiment the working temperature is 160° C., which generates saturated steam at 100 psia.

Power Collection Capacity

The total length of the GHB cylinders determines the power collection capacity of the system. In the current embodiment (without taking credit for the reflector panels) the diameter of the GHB cylinders is 0.5 meters and the total length of the cylinders is about 1200 meters, which provides a solar collection area of approximately 600 square meters. The addition of the reflector panel reduces the required total GHB length substantially from this value or provides some conservatism in the design.

Heat Sink

The preferred heat sink for the GBH is an inflatable radiating unit as described above.

Variations

Other Application of Radiating Unit

The radiator unit of the present invention will have numerous applications as a heat removal system for spacecraft. The radiator is configured as an inflatable structure, which when deflated can be rolled or folded into a lightweight, compact and rugged bundle. When in space the structure is expanded by inflation. Such applications include cooling large and dense bays of electronic components that are mounted to thermally conductive plates, which transfer their heat to a set of heat transfer fins through which hydrogen is circulated. Such electronics bays are common in satellite systems and the efficient transfer of heat from these interior spaces to external radiators is a continuing challenge.
Other systems can be envisioned that use the hydrogen in a radiator as a secondary loop, but not necessarily for use with a water-steam turbine system. A secondary heat transfer loop may be the preferred configuration where there is a potential that free molecular oxygen that may evolve in the electronics bays could react with the hydrogen in the presence of a spark. In systems where there is no danger of chemical reaction the hydrogen can be used as the primary heat transfer medium. For example, the hydrogen may be circulated directly through the electronics bays, and the hydrogen itself absorbs and carries heat away from the interior components to the external radiator system. In this case the interior surfaces within the bays will transfer heat directly to the gas.
Other systems can be envisioned which are hybrids of the hydrogen radiator with conventional heat transfer methods currently used in satellites, such as heat pipes. Heat pipes transfer heat very efficiently over short path lengths, but lose efficiency as the path lengths increase. A system of short heat pipes can be configured to absorb heat deep within the equipment bays and terminate in channels of flowing hydrogen which is part of the hydrogen radiator system. The terminated ends of such heat pipes could be fitted with fins to enhance the heat transfer to the hydrogen. The hydrogen would then carry the heat to the external radiator system.
In all applications, both primary and secondary, the external inflatable radiator system can be scaled in size to the application.

Gases Other than Hydrogen

Other applications can be envisioned in which another gas, such as nitrogen, methane, or any other suitable gas, replaces the hydrogen. This would reduce the heat transfer efficiency of the system due to the poorer thermal properties of most gasses. These gasses would also increase the overall mass somewhat, but would provide other benefits such as a lower permeability of the working fluid through the thin radiator surfaces or other components, and a lower potential for chemical reaction.
Other applications can be envisioned in which the radiator surfaces that are inflatable and made of coated Kapton (or other lightweight polymer) in the GHB system, is replaced in part by more rigid higher mass materials such as a lightweight metal or graphite which would impart greater strength to the system and reduce or eliminate gas loss problems.
Preferred embodiments of the present invention have been described in detail above. However, a great many variations from these specific embodiments could be made and will be obvious to persons skilled in the art to which this invention belongs. For example, as indicated in the above text, the size of the power generating system could be expanded almost without limit. Tubes could be of sizes different from the 0.5 meter diameter considering various criteria including those specifically referred to above. Various materials can be applied including some that are not even available today.
Therefore, the scope of the invention should be determined by the appended claims and their legal equivalence and not by the specific embodiments described above.

Claims

1. A radiator unit for radiating heat from a heat source into space, said radiating unit comprising:

A) an inflatable radiating panel comprising a plurality of inflatable tubes inflated with and adapted to contain and permit a circulation through the plurality of inflatable tubes of a low mass gas,

B) a sufficient quantity of low mass gas to inflate said plurality of inflatable tubes,

C) a blower unit for forcing the circulation of said low mass gas through the heat source and said plurality of inflatable tubes.

2. The radiator unit as in claim 1 wherein said low mass gas is hydrogen.

3. The radiator unit as in claim 1 wherein said unit is adapted to operate as a heat sink in a turbine generator system for the production of electric power in outer space.

4. The radiator unit as in claim 1 wherein said unit is adapted to operate as a heat sink in a thermoelectric system for the production of electric power in outer space.

5. The radiator unit as in claim 1 wherein said unit is adapted to operate as a heat sink for electrical equipment in outer space.

6. The radiator unit as in claim 1 wherein said radiator unit further comprises a heat exchanger for transfer of heat form a primary fluid to said low mass gas.

7. The radiator unit as in claim 2 wherein said plurality of inflatable tubes is comprised of a thermoset polyimide.

8. The radiator unit as in claim 2 wherein said plurality of inflatable tubes is comprised of a polyethylene terephthalate.

9. The radiator unit as in claim 4 wherein said plurality of inflatable tubes is lined with a metal foil.

10. The radiator unit as in claim 5 wherein said plurality of inflatable tubes is lined with a metal foil.

11. The radiator unit as in claim 1 wherein said low mass gas is a gas chosen from the following group of gasses: hydrogen, helium, nitrogen and neon.

12. The radiator unit as in claim 1 wherein said inflatable radiating panel is adapted for compaction into a small volume for transport into outer space and for inflation with said low mass gas transported into space in a high pressure gas bottle.

13. The radiator unit as in claim 3 wherein said electric generator system comprises:

A) a collector panel comprised of a plurality of collector tubes, each tube having walls comprised of thin flexible high-strength high temperature material with an outer surface having high absorptivity of solar radiation and low emissivity of thermal radiation,

B) a first working fluid defining a liquid phase and a vapor phase,

C) a turbine generator system adapted to electrical power from the vapor phase of the working fluid,

14. The system as in claim 13 wherein the first working fluid is water.

15. The system as in claim 13 wherein the walls of the collector tubes are comprised of a composite material comprised of a polymer fiber.

16. The system as in claim 13 wherein the entire system is adapted to fit in a cargo space of a space vehicle.

17. The system as in claim 16 wherein the space vehicle is an Atlas 5 HLV rocket.