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Lightweight radiator fin made of tapered panels bent together at ends and joined to heat pipe in middle. Taper chosen for minimum mass per unit of radiated power. Intended primarily for use as part of cooling system of space station, also used on Earth to cool equipment in laboratory vacuum. Possible to modify basic design equations to design minimum-mass fins for cooling by conduction or convection.
Experimental results of tests made at the Langley Memorial Aeronautical Laboratory are presented to show how heat-transfer coefficients can he increased by a method utilizing the high rate of heat transfer known to exist on any heat-transfer surface in the region adjacent to the edge on which the cooling or heating fluid impinges. The results show that, for the same pressure drop, the average surface heat-transfer.coefficient can be increased 50 to 100 percent when a cooling surface having a length of four inches in the direction of fluid flow is cut to form twenty fins with a length of 0.2 inch in the direction of fluid flow and the fins are sharpened and staggered in the air stream. The percentage of increase in the surface heat-transfer coefficient obtained as a result of shortening the length of the cooling surface varies with the pressure drop of the cooling fluid in passing the surface, the increase being largest when small pressure drop is used and smallest when high pressure drop is used.
A high efficiency forced air heat sink assembly employs a split feed transverse flow configuration to minimize the length of the air flow path through at least two separated fin structures. Different embodiments use different fin structure material configurations including honeycomb, corrugated and serpentine. Each such embodiment uses a thermally conductive plate having opposed exterior surfaces; one for receiving a component to be cooled and one for receiving the fin structures. The serpentine structured fin embodiment employs a plurality of fin supports extending from the plate and forming a plurality of channels for receiving the fin structures. A high thermal conductivity bondant, such as metal-filled epoxy, may be used to bond the fin structures to either the plate or the fin supports. Dip brazing and soldering may also be employed depending upon the materials selected.
This paper describes the thermal analysis, design, and testing of a dedicated cooling system for a Spartan spacecraft payload. A simple reliable design that requires minimum power consumption and minimum weight was developed. The payload has a CCD detector that must be maintained at a temperature of approximately -40 C or colder. The cooling system consists of a fin radiator, dual redundant heat pipes, and a thermal electric device (TED). The system was analytically modeled through the use of the Simplified Shuttle Payload Thermal Analyzer (SSPTA) computer program. A thermal test of the system simulating flight conditions was conducted to correlate the computer model and verify performance specifications.
After fueling and prior to launch, the Advanced Stirling Radioisotope Generator (ASRG) will be stored for a period of time then moved to the launch pad for integration with the space probe and mounting on the launch vehicle. During this time, which could be as long as 3 years, the ASRG will operate continuously with heat rejected from the housing and fins. Typically, the generator will be cooled by forced convection using fans. During some of the ground operations, maintaining forced convection may add significant complexity, so allowing natural convection may simplify operations. A test was conducted on the ASRG Engineering Unit (EU) to quantify temperatures and operating parameters with natural convection only and determine if the EU could be safely operated in such an environment. The results show that with natural convection cooling the ASRG EU Stirling convertor pressure vessel temperatures and other parameters had significant margins while the EU was operated for several days in this configuration. Additionally, an update is provided on ASRG EU testing at NASA Glenn Research Center, where the ASRG EU has operated for over 16,000 hr and underwent extensive testing.
An investigation to determine and correlate the experimental surface heat-transfer coefficients of finned cylinders with different air-stream cooling arrangements was conducted at the Langley Memorial Aeronautical Laboratory from 1932 to 1938. The investigation covered the determination of the effect of fin width, fin space, fin thickness, and cylinder diameter on the heat transfer. Wind-tunnel tests were made in the free air stream with and without baffles and also with various devices for creating a turbulent air stream. Tests were also made with blower.
Three joining methods were considered for use in fabricating cooling coils for the National Transonic Facility. After analysis and preliminary testing, soldering was chosen as the cooling coil joining technique over mechanical force fit and brazing techniques. Charpy V-Notch tests, cyclic thermal tests (ambient to 77.8 K) and tensile tests at cryogenic temperatures were performed on solder joints to evaluate their structural integrity. It was determined that low temperature solder can be used to ensure good fin-to-tube contact for cooling-coil applications.
This program covered the design, fabrication and testing of an advanced development model uninterrupted power system. The input and output requirements imposed on the power processor were specified such that the unit is electrically interchangeable with existing power systems used by the Federal Aviation Administration in installations which have a history of failure due to electrical transient conditions. Input power is from either of two single-phase ac power sources or batteries with electronic selection and transfer between power sources. Battery reconditioning is automatic when either ac source is present. The output power is rated at 84OW; the nominal output is 24V at 35A. Within the 84OW limit, the regulated output voltage is adjustable from 22V to 30Vdc. Protection against continuous overloading or short circuit is provided. The unit is packaged in a standard 19-inch rack mount configuration with 7-inch panel height. Controls are on the front panel with power input and output through connectors on the rear surface. Cooling is by free convection from fin areas located on the side and rear panels. The packaged unit weighs 52.8 lbs., which can be reduced significantly if a three-phase ac power source is used.
Selected literature on heat-transfer and pressure losses for airflow through passages for several cooling methods generally applicable to gas turbine blades is reviewed. Some useful correlating equations are highlighted. The status of turbine-blade internal air-cooling technology for both nonrotating and rotating blades is discussed and the areas where further research is needed are indicated. The cooling methods considered include convection cooling in passages, impingement cooling at the leading edge and at the midchord, and convection cooling in passages, augmented by pin fins and the use of roughened internal walls.
With the aid of a heating device, the heat transfer to cylinders with conical fins of various forms is determined both for shrouded and exposed cylinders. Simultaneously the pressure drop for overcoming the resistance to the motion of air between the fins of the enclosed cylinder is measured. Thus the relations between the heat transfer and the energy required for cooling are discovered. The investigations show that the heat transfer in a conducted air flow is much greater than in a free current and that further improvement, as compared with free exposure, is possible through narrower spaces between the fins.
Measurements were made of the cooling on the fronts of model cylinders in a conventional cowling for cooling in both the ground and the cruising conditions. The mechanisms of front and rear cooling are essentially different. Cooling on the rear baffled part of the cylinders continually increases with increasing fin width. For the front of the cylinder, an optimum fin width was found to exist beyond which an increase in width reduced the heat transfer. The heat transfer coefficient on the front of the cylinders was larger on the side of the cylinder facing the propeller swirl than on the opposite side. This effect became more pronounced as the fin width was increased. These results are introductory to the study of front cooling and show the general effect of several test parameters.
A heat exchanger, as exemplified by a rocket combustion chamber, is constructed by stacking thin metal rings having microsized openings therein at selective locations to form cooling passages defined by an inner wall, an outer wall and fins. Suitable manifolds are provided at each end of the rocket chamber. In addition to the cooling channel openings, coolant feed openings may be formed in each of rings. The coolant feed openings may be nested or positioned within generally U-shaped cooling channel openings. Compression on the stacked rings may be maintained by welds or the like or by bolts extending through the stacked rings.
Prototype adsorption device used, for example, in adsorption heat pump, to store natural gas to power automobile, or to separate components of fluid mixtures. Device includes activated carbon held together by binder and molded into finned heat-transfer device providing rapid heating or cooling to enable rapid adsorption or desorption of fluids. Concepts of design and fabrication of device equally valid for such other highly thermally conductive devices as copper-finned tubes, and for such other high-surface-area sorbents as zeolites or silicates.
The paper describes four separate advanced space radiator concepts that have been pursued in an integrated effort to develop multipayload use, low-cost heat rejection systems which can overcome the limitations of current radar systems. These four concepts are the following: (1) a wide heat load range, modularized space radiation system; (2) a heat rejection subsystem designed as a compact self-contained heat rejection module which provides a refrigeration cycle to accomodate a wide variation in payload heat loads and cooling temperature requirements; (3) a lightweight flexible fin radiator system with heat rejection capability; and (4) a radiator which does not require a circulating coolant on the radiator panel with heat pipes and thus particularly applies to very-long-duration free-flying payloads where long life reliability is an overriding design parameter. The Orbiter active thermal control system is briefly described.
The heat-transfer coefficients have been determined for five steel cylinders having fins 1.22 inches wide and the spacing between the fins ranging from 0.022 to 0.131 inch. The cylinders were tested with and without baffles in a wind tunnel; they were also tested enclosed in jackets with the cooling air supplied by a blower. A maximum heat transfer was reached at a fin space of about 0.45 inch for the cylinders tested with each of the three methods of cooling investigated. The rise in temperature of the air passing between the fins and the change in flow pattern were found to be important factors limiting the heat transfer that may be obtained by decreasing the fin space. The use of baffles for directing the air around the cylinders with closely spaced fins proved very effective in increasing the over-all heat-transfer coefficient, provided that the spacing was not appreciably less than that for maximum heat transfer.
This report gives the results of an experimental determination of the temperature distribution in and the heat dissipation from a cylindrical finned surface for various fin-plane/air-stream angles. A steel cylinder 4.5 inches in diameter having slightly tapered fins of 0.30-inch pitch and 0.6 -inch width was equipped with an electrical heating unit furnishing 13 to 248 B.T.U. per hour per square inch of inside wall area. Air at speeds form 30 to 150 miles per hour was directed at seven different angles from 0 degrees to 90 degrees with respect to the fin planes. The tests show the best angle for cooling at all air speeds to be about 45 degrees. With the same temperature for the two conditions and with an air speed of 76 miles per hour, the heat input to the cylinder can be increased 50 percent at 45 degrees fin-plane/air-stream angle over that at 0 degrees.
A generalized one-dimensional computer code for analyzing the flow and heat transfer in the turbomachinery cooling passages was developed. This code is capable of handling rotating cooling passages with turbulators, 180 degree turns, pin fins, finned passages, by-pass flows, tip cap impingement flows, and flow branching. The code is an extension of a one-dimensional code developed by P. Meitner. In the subject code, correlations for both heat transfer coefficient and pressure loss computations were developed to model each of the above mentioned type of coolant passages. The code has the capability of independently computing the friction factor and heat transfer coefficient on each side of a rectangular passage. Either the mass flow at the inlet to the channel or the exit plane pressure can be specified. For a specified inlet total temperature, inlet total pressure, and exit static pressure, the code computers the flow rates through the main branch and the subbranches, flow through tip cap for impingement cooling, in addition to computing the coolant pressure, temperature, and heat transfer coefficient distribution in each coolant flow branch. Predictions from the subject code for both nonrotating and rotating passages agree well with experimental data. The code was used to analyze the cooling passage of a research cooled radial rotor.
A generalized one-dimensional computer code for analyzing the flow and heat transfer in the turbomachinery cooling passages was developed. This code is capable of handling rotating cooling passages with turbulators, 180 degree turns, pin fins, finned passages, by-pass flows, tip cap impingement flows, and flow branching. The code is an extension of a one-dimensional code developed by P. Meitner. In the subject code, correlations for both heat transfer coefficient and pressure loss computations were developed to model each of the above mentioned type of coolant passages. The code has the capability of independently computing the friction factor and heat transfer coefficient on each side of a rectangular passage. Either the mass flow at the inlet to the channel or the exit plane pressure can be specified. For a specified inlet total temperature, inlet total pressure, and exit static pressure, the code computers the flow rates through the main branch and the subbranches, flow through tip cap for impingement cooling, in addition to computing the coolant pressure, temperature, and heat transfer coefficient distribution in each coolant flow branch. Predictions from the subject code for both nonrotating and rotating passages agree well with experimental data. The code was used to analyze the cooling passage of a research cooled radial rotor.