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At least 37 records · Page 2

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NASA NTRS1939

Correction of Temperatures of Air-Cooled Engine Cylinders for Variation in Engine and Cooling Conditions

Factors are obtained from semiempirical equations for correcting engine-cylinder temperatures for variation in important engine and cooling conditions. The variation of engine temperatures with atmospheric temperature is treated in detail, and correction factors are obtained for various flight and test conditions, such as climb at constant indicated air speed, level flight, ground running, take-off, constant speed of cooling air, and constant mass flow of cooling air. Seven conventional air-cooled engine cylinders enclosed in jackets and cooled by a blower were tested to determine the effect of cooling-air temperature and carburetor-air temperature on cylinder temperatures. The cooling air temperature was varied from approximately 80 degrees F. to 230 degrees F. and the carburetor-air temperature from approximately 40 degrees F. to 160 degrees F. Tests were made over a large range of engine speeds, brake mean effective pressures, and pressure drops across the cylinder. The correction factors obtained experimentally are compared with those obtained from the semiempirical equations and a fair agreement is noted.

Schey, Oscar W↗

NASA NTRS1941

The problem of cooling an air-cooled cylinder on an aircraft engine

An analysis of the cooling problem has been to show by what means the cooling of an air-cooled aircraft engine may be improved. Each means of improving cooling is analyzed on the basis of effectiveness in cooling with respect to power for cooling. The altitude problem is analyzed for both supercharged and unsupercharged engines. The case of ground cooling is also discussed. The heat-transfer process from the hot gases to the cylinder wall is discussed on the basis of the fundamentals of heat transfer and thermodynamics. Adiabatic air-temperature rise at a stagnation point in compressible flow is shown to depend only on the velocity of flow.

Brevoort, M J↗

NASA NTRS1959

Air-Cooled Turbine Blades with Tip Cap For Improved Leading-Edge Cooling

An investigation was conducted in a modified turbojet engine to determine the cooling characteristics of the semistrut corrugated air- cooled turbine blade and to compare and evaluate a leading-edge tip cap as a means for improving the leading-edge cooling characteristics of cooled turbine blades. Temperature data were obtained from uncapped air-cooled blades (blade A), cooled blades with the leading-edge tip area capped (blade B), and blades with slanted corrugations in addition to leading-edge tip caps (blade C). All data are for rated engine speed and turbine-inlet temperature (1660 F). A comparison of temperature data from blades A and B showed a leading-edge temperature reduction of about 130 F that could be attributed to the use of tip caps. Even better leading-edge cooling was obtained with blade C. Blade C also operated with the smallest chordwise temperature gradients of the blades tested, but tip-capped blade B operated with the lowest average chordwise temperature. According to a correlation of the experimental data, all three blade types 0 could operate satisfactorily with a turbine-inlet temperature of 2000 F and a coolant flow of 3 percent of engine mass flow or less, with an average chordwise temperature limit of 1400 F. Within the range of coolant flows investigated, however, only blade C could maintain a leading-edge temperature of 1400 F for a turbine-inlet temperature of 2000 F.

Calvert, Howard F.↗

NASA NTRS2015

Analysis of Turbine Blade Relative Cooling Flow Factor Used in the Subroutine Coolit Based on Film Cooling Correlations

Heat transfer correlations of data on flat plates are used to explore the parameters in the Coolit program used for calculating the quantity of cooling air for controlling turbine blade temperature. Correlations for both convection and film cooling are explored for their relevance to predicting blade temperature as a function of a total cooling flow which is split between external film and internal convection flows. Similar trends to those in Coolit are predicted as a function of the percent of the total cooling flow that is in the film. The exceptions are that no film or 100 percent convection is predicted to not be able to control blade temperature, while leaving less than 25 percent of the cooling flow in the convection path results in nearing a limit on convection cooling as predicted by a thermal effectiveness parameter not presently used in Coolit.

Air cooling↗

NASA NTRS2018

Three Stage Cool Flame Droplet Burning Behavior of n-Alkane Droplets at Elevated Pressure Conditions: Hot, Warm and Cool Flame

Transient, isolated n-alkane droplet combustion is simulated at elevated pressure for helium-diluent substituted-air mixtures. We report the presence of unique quasi-steady, three-stage burning behavior of large sphero-symmetric n-alkane droplets at these elevated pressures and helium substituted ambient fractions. Upon initiation of reaction, hot-flame diffusive burning of large droplets is initiated that radiatively extinguishes to establish cool flame burning conditions in nitrogen/oxygen “air” at atmospheric and elevated pressures. However, at elevated pressure and moderate helium substitution for nitrogen ( X He > 20%), the initiated cool flame burning proceeds through two distinct, quasi-steady-state, cool flame burning conditions. The classical “Hot flame ” ( ∼1500 K) radiatively extinguishes into a “Warm flame ” burning mode at a moderate maximum reaction zone temperature ( ∼ 970 K), followed by a transition to a lower temperature ( ∼765 K), quasi-steady “Cool flame ” burning condition. The reaction zone (“flame”) temperatures are associated with distinctly different yields in intermediate reaction products within the reaction zones and surrounding near-field, and the flame-standoff ratios characterizing each burning mode progressively decrease. The presence of all three stages first appears with helium substitution near 20%, and the duration of each stage is observed to be strongly dependent on helium substitutions level between 20–60%. For helium substitution greater than 60%, the hot flame extinction is followed by only the lower temperature cool flame burning mode. In addition to the strong coupling between the diffusive loss of both energy and species and the slowly evolving degenerate branching in the low and negative temperature coefficient (NTC) kinetic regimes, the competition between the low-temperature chain branching and intermediate-temperature chain termination reactions control the “Warm” and “Cool ” flame quasi-steady conditions and transitioning dynamics.

Droplet Combustion↗

NASA NTRS1972

Comparison of heat transfer characteristics of three cooling configurations for air-cooled turbine vanes tested in a turbojet engine

A comparison was made of the heat transfer characteristics of three air cooled vanes. The vanes incorporated cooling schemes such as impingement cooling, film cooling, and convection cooling with and without extended surfaces. A redesign study was made for two vanes to improve the cooling effectiveness. An average impingement heat transfer coefficient was calculated on the bases of experimentally determined temperatures at the leading edge and a one dimensional heat transfer calculation. This heat transfer coefficient was compared with existing impingement heat transfer correlations.

Yeh, F. C.↗

NASA NTRS1974

Comparison of cooling effectiveness of turbine vanes with and without film cooling

The cooling effectiveness of three film-cooled vanes were compared to the cooling effectiveness of two non-film-cooled vanes. The comparison indicated that, for the vane configurations and test conditions examined, film cooling had an adverse effect near the suction-surface trailing edge of the vanes. Film cooling was found to be beneficial on the pressure surface of the vanes.

Yeh, F. C.↗

NASA NTRS1996

Cooling Duct Analysis for Transpiration/Film Cooled Liquid Propellant Rocket Engines

The development of a low cost space transportation system requires that the propulsion system be reusable, have long life, with good performance and use low cost propellants. Improved performance can be achieved by operating the engine at higher pressure and temperature levels than previous designs. Increasing the chamber pressure and temperature, however, will increase wall heating rates. This necessitates the need for active cooling methods such as film cooling or transpiration cooling. But active cooling can reduce the net thrust of the engine and add considerably to the design complexity. Recently, a metal drawing process has been patented where it is possible to fabricate plates with very small holes with high uniformity with a closely specified porosity. Such a metal plate could be used for an inexpensive transpiration/film cooled liner to meet the demands of advanced reusable rocket engines, if coolant mass flow rates could be controlled to satisfy wall cooling requirements and performance. The present study investigates the possibility of controlling the coolant mass flow rate through the porous material by simple non-active fluid dynamic means. The coolant will be supplied to the porous material by series of constant geometry slots machined on the exterior of the engine.

Micklow, Gerald J.↗

NASA NTRS1981

Solute partitioning under continuous cooling conditions as a cooling rate indicator

A model of solute partitioning in a finite body under conditions of continuous cooling is developed for the determination of cooling rates from concentration profile data, and applied to the partitioning of zirconium between ilmenite and ulvospinel in the Apollo 15 Elbow Crater rocks. Partitioning in a layered composite solid is described numerically in terms of concentration profiles and diffusion coefficients which are functions of time and temperature, respectively; a program based on the model can be used to calculate concentration profiles for various assumed cooling rates given the diffusion coefficients in the two phases and the equilibrium partitioning ratio over a range of temperatures. In the case of the Elbow Rock gabbros, the cooling rates are calculated from measured concentration ratios 10 microns from the interphase boundaries under the assumptions of uniform and equilibrium initial conditions at various starting temperatures. It is shown that the specimens could not have had uniform concentrations profiles at the previously suggested initial temperature of 1350 K. It is concluded that even under conditions where the initial temperature, grain sizes and solute diffusion coefficients are not well characterized, the model can be used to estimate the cooling rate of a grain assemblage to within an order of magnitude.

Onorato, P. I. K.↗

NASA NTRS1990

The utilization of an infrared imaging system as a cooling slot blockage detector in the inspection of a transpiration cooled nozzle

A comprehensive examination of the 8 foot temperature tunnel's transpiration cooled nozzle was completed using an infrared imaging radiometer to locate regions of cooling flow irregularities caused by obstruction of three or more adjacent cooling slots. Restrictions in the cooling flow were found and cataloged. Blockages found were due primarily to the presence of residual phosphoric acid being discharged from some of the cooling slots. This acid was used during construction of the nozzle components and was to have been purged prior to its delivery to the NASA Langley Research Center (LaRC). In addition, a radial displacement of one selection of discs located in the spool piece was inspected and cataloged for future reference. There did not seem to be a serious restriction of flow in this defect, but evidence from the infrared images indicated reduced slot activity within the gouge. The radiometer survey uncovered regions where closer inspection is recommended but did not cover the entire surface area of the three nozzle subsections due to equipment limitations. A list of areas with suspected problems is included in Appendix A.

Borg, Stephen E.↗

NASA NTRS1930

Drag and Cooling with Various Forms of Cowling for a "Whirlwind" Radial Air-Cooled Engine I

This report presents the results of an investigation undertaken in the 20-foot Propeller Research Tunnel at Langley Field on the cowling of radial air-cooled engines. A portion of the investigation has been completed, in which several forms and degrees of cowling were tested on Wright "Whirlwind" J-5 engine mounted in the nose of a cabin fuselage. The cowlings varied from the one extreme of an entirely exposed engine to the other in which the engine was entirely inclosed. Cooling tests were made and each cowling modified, if necessary, until the engine cooled approximately as satisfactorily as when it was entirely exposed. Drag tests were then made with each form of cowling, and the effect of the cowling on the propulsive efficiency determined with a metal propeller. The propulsive efficiency was found to be practically the same with all forms of cowling. The drag of the cabin fuselage with uncowled engine was found to be more than three times as great as the drag of the fuselage with engine removed and nose rounded. The conventional forms of cowling, in which at least the tops of the cylinder heads and valve gear are exposed, reduce the drag somewhat, but the cowling entirely covering the engine reduces it 2.6 times as much as the best conventional one. The decrease in drag due to the use of spinners proved to be almost negligible. The use of the cowling completely covering the engine seems entirely practical as regards both cooling and maintenance under service conditions. It must be carefully designed, however, to cool properly. With cabin fuselages its use should result in a substantial increase in high speed over that obtained with present forms of cowling on engines similar in contour to the J-5. (author)

Weick, Fred E↗

NASA NTRS1939

Cooling on the Front of an Air-cooled Engine Cylinder in a Conventional Engine Cowling

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.

Brevoort, M J↗

NASA NTRS1945

Correlation of the Characteristics of Single-Cylinder and Flight Engines in Tests of High-Performance Fuels in an Air-Cooled Engine I : Cooling Characteristics

Variable charge-air flow, cooling-air pressure drop, and fuel-air ration investigations were conducted to determine the cooling characteristics of a full-scale air-cooled single cylinder on a CUE setup. The data are compared with similar data that were available for the same model multicylinder engine tested in flight in a four-engine airplane. The cylinder-head cooling correlations were the same for both the single-cylinder and the flight engine. The cooling correlations for the barrels differed slightly in that the barrel of the single-cylinder engine runs cooler than the barrel of te flight engine for the same head temperatures and engine conditions.

Wilson, Robert W.↗

NASA NTRS1945

Cooling of Gas Turbines: Effectiveness of Rim Cooling of Blades - 2

An analysis of rim cooling, which cools the blade by condition alone, was conducted. Gas temperatures ranged from 1300 degrees to 1900 degrees F and rim temperatures from 0 degrees to 1000 degrees F below gas temperatures. Results show that gas temperature increases up to 200 degrees F are permissible provided that the blades are cooled by 400 degrees to 500 degrees F below the gas temperature. Relatively small amounts of blade cooling, at constant gas temperature, give large increases in blade life. Dependence of rim cooling on heat-transfer coefficient, blade dimensions, and thermal conductivity is determined by a single parameter.

ATI COLLECTION↗

NASA NTRS1975

A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Active cooling system analysis

The effects of fuselage cross section and structural arrangement on the performance of actively cooled hypersonic cruise vehicles are investigated. An active cooling system which maintains the aircraft's entire surface area at temperatures below 394 K at Mach 6 is developed along with a hydrogen fuel tankage thermal protection system. Thermodynamic characteristics of the actively cooled thermal protection systems established are summarized. Design heat loads and coolant flowrate requirements are defined for each major structural section and for the total system. Cooling system weights are summarized at the major component level. Conclusions and recommendations are included.

Stone, J. E.↗

NASA NTRS1946

Analysis of cooling limitations and effect of engine-cooling improvements on level-flight cruising performance of four-engine heavy bomber

The NACA has developed means, including an injection impeller and ducted head baffles, to improve the cooling characteristics of the 3350-cubic-inch-displacement radial engines installed in a four-engine heavy bomber. The improvements afforded proper cooling of the rear-row exhaust-valve seats for a wide range of cowl-flap angles, mixture strengths, and airplane speeds. The results of flight tests with this airplane are used as a basis for a study to determine the manner and the extent to which the airplane performance was limited by engine cooling. By means of this analysis for both the standard airplane and the airplane with engine-cooling modifications, comparison of the specific range at particular conditions and comparison of the cruising-performance limitations was made.

Marble, Frank E↗

NASA NTRS1993

Cation ordering in orthopyroxenes and cooling rates of meteorites: Low temperature cooling rates of Estherville, Bondoc and Shaw

The cooling rates of meteorites provide important constraints on the size of their parent bodies, and their accretionary and evolutionary histories. However, the cooling rates obtained so far from the commonly used metallographic, radiometric and fission-track methods have been sometimes quite controversial, such as in the case of the mesosiderites and the meteorite Shaw. We have undertaken a systematic study of the cooling rates of meteorites using a different approach, which involves single crystal x-ray determination of Fe(2+)-Mg ordering in orthopyroxenes (OP(x)) in meteorites, subject to bulk compositional constraints, and numerical simulation of the evolution of the ordering state as a function of cooling rate, within the framework of the thermodynamic and kinetic principles governing cation ordering. We report the results obtained for OP(x) crystals from Shaw and two mesosiderites, Estherville and Bondoc.

Ganguly, J.↗

NASA NTRS1994

The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

Kacynski, Kenneth J.↗