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At least 397 records · Page 22

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NASA NTRS1975

Aerodynamic performance of a fully film cooled core turbine vane tested with cold air in a two-dimensional cascade

The aerodynamic performance of a fully film cooled core turbine vane was investigated experimentally in a two-dimensional cascade of 10 vanes. Three of the 10 vanes were cooled; the others were solid (uncooled) vanes. Cold air was used for both the primary and coolant flows. The cascade test covered a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95 and a range of coolant flow rates to 7.5 percent of the primary flow. The coolant flow was varied by changing the coolant supply pressure. The principal measurements were cross-channel surveys of exit total pressure, static pressure, and flow angle. The results presented include exit survey data and overall performance in terms of loss, flow angle, and weight flow for the range of exit velocity ratios and coolant flows investigated. The performance of the cooled vane is compared with the performance of an uncooled vane of the same profile and also with the performance obtained with a single cooled vane in the 10-vane cascade.

Stabe, R. G.↗

NASA NTRS1976

Turbine vane gas film cooling with injection in the leading edge region from a single row of spanwise angled holes

An experimental study of gas film cooling was conducted on a 3X size model turbine vane. Injection in the leading edge region was from a single row of holes angled in a spanwise direction. Measurements of the local heat flux downstream from the row of coolant holes, both with and without film coolant flow, were used to determine the film cooling performance presented in terms of the Stanton number ratio. Results for a range of coolant blowing ratio, M = 0 to 2.0, indicate a reduction in heat flux of up to 15 to 30 percent at a point 10 to 11 hole diameters downstream from injection. An optimum coolant blowing ratio corresponds to a coolant-to-freestream velocity ratio in the range of 0.5. The shallow injection angle resulted in superior cooling performance for injection closest to stagnation, while the effect of injection angle was insignificant for injection further from stagnation.

Lecuyer, M. R.↗

NASA NTRS1976

Rocket engine thrust chamber wall temperature distribution calculation and analysis

An analytical computational concept is presented which predicts the temperature profiles along a regeneratively cooled thrust chamber wall on the hot gas side and on the coolant side, and also the coolant bulk temperature profile. The computational model is based upon a coupling of the boundary layer heat transfer process with the heat transfer process through the chamber wall and the coolant flow heat absorption. The calculation is started with approximate temperature distributions for the hot gas side wall and the coolant flow. The iteration process of the computer program is terminated when the total heat transfer rates from the hot gas boundary layer to the wall and from the wall to the coolant are equal. The computer program for the integration of regenerative cooling process to a thrust chamber is kept general such that this program can be used with any boundary layer analysis computer program for temperature profile and heat transfer studies. A sample application of this concept is shown by using a boundary layer analysis program for the RL10 rocket engine thrust chamber.

Saha, H.↗

NASA NTRS1976

Two-dimensional cold-air cascade study of a film-cooled turbine stator blade. 3: Effect of hole size on single-row and multirow ejection

The effect of coolant discharge on the aerodynamic performance of a film cooled turbine stator blade was determined. The blade had the same number, location, and injection angle of coolant holes, but the coolant hole diameters were one half that of a previously investigated blade. Otherwise the blades were the same. Tests with discharge from individual coolant rows and multiple coolant rows, including full film discharge are studied. The results of the blade with smaller holes are reported and compared with the blades with larger holes.

Prust, H. W., Jr.↗

NASA NTRS1981

Cooling system for high speed aircraft

The system eliminates the necessity of shielding an aircraft airframe constructed of material such as aluminum. Cooling is accomplished by passing a coolant through the aircraft airframe, the coolant acting as a carrier to remove heat from the airframe. The coolant is circulated through a heat pump and a heat exchanger which together extract essentially all of the added heat from the coolant. The heat is transferred to the aircraft fuel system via the heat exchanger and the heat pump. The heat extracted from the coolant is utilized to power the heat pump. The heat pump has associated therewith power turbine mechanism which is also driven by the extracted heat. The power turbines are utilized to drive various aircraft subsystems, the compressor of the heat pump, and provide engine cooling.

Lawing, P. L.↗

NASA NTRS1983

Aerodynamic effect of a honeycomb rotor tip shroud on a 50.8-centimeter-tip-diameter core turbine

A 50.8-cm-tip-diameter turbine equipped with a rotor tip shroud of hexagonal cell (or honeycomb) cross section has been tested in warm air (416 K) for a range of shroud coolant to primary flow rates. Test results were also obtained for the same turbine operated with a solid shroud for comparison. The results showed that the combined effect of the honeycomb shroud and the coolant flow was to cause a reduction of 2.8 points in efficiency at design speed, pressure ratio, and coolant flow rate. With the coolant system inactivated, the honeycomb shroud caused a decrease in efficiency of 2.3 points. These results and those obtained from a small reference turbine indicate that the dominant factor governing honeycomb tip shroud loss is the ratio of honeycomb depth to blade span. The loss results of the two shrouds could be correlated on this basis. The same honeycomb and coolant effects are expected to occur for the hot (2200 K) version of this turbine.

Moffitt, T. P.↗

NASA NTRS1985

Jet pump-drive system for heat removal

A jet pump, in combination with a TEMP, is employed to assure safe cooling of a nuclear reactor after shutdown. A TEMP, responsive to the heat from the coolant in the secondary flow path, automatically pumps the withdrawn coolant to a higher pressure and thus higher velocity compared to the main flow. The high velocity coolant is applied as a driver flow for the jet pump which has a main flow chamber located in the main flow circulation pump. Upon nuclear shutdown and loss of power for the main reactor pumping system, the TEMP/jet pump combination continues to boost the coolant flow in the direction it is already circulating. During the decay time for the nuclear reactor, the jet pump keeps running until the coolant temperature drops to a lower and safe temperature. At this lower temperature, the TEMP/jet jump combination ceases its circulation boosting operation. The TEMP/jet pump combination is automatic, self-regulating and provides an emergency pumping system free of moving parts.

French, J. R.↗

NASA NTRS1990

Shock/shock interference on a transpiration cooled hemispherical model

Experimental results are presented which show the effectiveness of transpiration cooling in reducing the peak heat flux caused by an impinging shock on a bow shock of a hemispherical model. The 12-inch diameter hemispherical transpiration model with helium coolant was tested in the Calspan 48-inch Hypersonic Shock Tunnel at nominal Mach 12.1 and freestream unit Reynolds number of 0.33 x 10 to the 6th/ft. An incident shock wave, generated by a blunt flat-plate shock generator inclined at 10 deg to the freestream, intersected the bow shock of the model to produce shock/shock interference. The stagnation heat flux without coolant or shock/shock interference was about 1.6 times a smooth surface laminar prediction due to effective roughness of the coolant ejection slots. A coolant mass flux 31 percent of the freestream mass flux reduced the stagnation heat flux to zero without shock/shock interference. However, for the same coolant mass flux and with shock/shock interference the peak heat flux was only reduced 8.3 percent, even though the total integrated heat load was reduced.

Nowak, Robert J.↗

NASA NTRS1992

An experimental investigation of high-aspect-ratio cooling passages

An experimental investigation was conducted to evaluate the effectiveness of using high-aspect-ratio cooling passages to improve the life and reduce the coolant pressure drop in high-pressure rocket thrust chambers. A plug-nozzle rocket-engine test apparatus was used to test two cylindrical chambers with low-aspect-ratio cooling passages and one with high-aspect-ratio cooling passages. The chambers were cyclically tested and data were taken over a wide range of coolant mass flows. The results showed that for the same coolant pressure drop, the hot-gas-side wall temperature of the high-aspect-ratio chamber was 30 percent lower than the baseline low-aspect-ratio chamber, resulting in no fatigue damage to the wall. The coolant pressure drop for the high-aspect-ratio chamber was reduced in increments to one-half that of the baseline chamber, by reducing the coolant mass flow, and still resulted in a reduction in the hot-gas-side wall temperature when compared to the low-aspect-ratio chambers.

Carlile, Julie A.↗

NASA NTRS1952

Internal-liquid-film-cooling Experiments with Air-stream Temperatures to 2000 Degrees F. in 2- and 4-inch-diameter Horizontal Tubes

Report presents the results of an investigation conducted to determine the effectiveness of liquid-cooling films on the inner surfaces of tubes containing flowing hot air. Experiments were made in 2- and 4-inch-diameter straight metal tubes with air flows at temperatures from 600 degrees to 2000 degrees F. and diameter Reynolds numbers from 2.2 to 14 x 10(5). The film coolant, water, was injected around the circumference at a single axial position on the tubes at flow rates from 0.02 to .24 pound per second per foot of tube circumference (0.8 to 12 percent of the air flow). Liquid-coolant films were established and maintained around and along the tube wall in concurrent flow with the hot air. The results indicated that, in order to film cool a given surface area with as little coolant flow as possible, it may be necessary to limit the flow of coolant introduced at a single axial position and to introduce it at several axial positions. The flow rate of inert coolant required to maintain liquid-film cooling over a given area of tube surface can be estimated when the gas-flow conditions are known by means of a generalized plot of the film-cooling data.

Kinney, George R↗

NASA NTRS1995

Leading edge film cooling effects on turbine blade heat transfer

An existing three dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of spanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. It is found that with the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P/d= 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface due to coolant jet lift-off. For P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dimensional.

Garg, Vijay K.↗

NASA NTRS2000

Unsteady High Turbulence Effects on Turbine Blade Film Cooling Heat Transfer Performance Using a Transient Liquid Crystal Technique

Unsteady wake effect, with and without trailing edge ejection, on detailed heat transfer coefficient and film cooling effectiveness distributions is presented for a downstream film-cooled gas turbine blade. Tests were performed on a five-blade linear cascade at an exit Reynolds number of 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.4 to 1.2; air and CO2 were used as coolants to simulate different density ratios. Surface heat transfer and film effectiveness distributions were obtained using a transient liquid crystal technique; coolant temperature profiles were determined with a cold wire technique. Results show that Nusselt numbers for a film cooled blade are much higher compared to a blade without film injection. Unsteady wake slightly enhances Nusselt numbers but significantly reduces film effectiveness versus no wake cases. Nusselt numbers increase only slic,htly but film cooling, effectiveness increases significantly with increasing, blowing ratio. Higher density coolant (CO2) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher 0 effectiveness at lower blowing ratios (M = 0.8). Trailing edge ejection generally has more effect on film effectiveness than on the heat transfer, typically reducing film effectiveness and enhancing heat transfer. Similar data is also presented for a film cooled cylindrical leading edge model.

Han, J. C.↗

NASA NTRS1947

Cooling of Gas Turbines: Analysis of Rotor and Blade Temperatures in Liquid-Cooled Gas Turbines - 3

A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Brown, W. Byron↗

NASA NTRS2003

Membrane-Based Gas Traps for Ammonia, Freon-21, and Water Systems to Simplify Ground Processing

Gas traps are critical for the smooth operation of coolant loops because gas bubbles can cause loss of centrifugal pump prime, interference with sensor readings, inhibition of heat transfer, and blockage of passages to remote systems. Coolant loops are ubiquitous in space flight hardware, and thus there is a great need for this technology. Conventional gas traps will not function in micro-gravity due to the absence of buoyancy forces. Therefore, clever designs that make use of adhesion and momentum are required for adequate separation, preferable in a single pass. The gas traps currently used in water coolant loops on the International Space Station are composed of membrane tube sets in a shell. Each tube set is composed of a hydrophilic membrane (used for water transport and capture of bubbles) and a hydrophobic membrane (used for venting of air bubbles). For the hydrophilic membrane, there are two critical pressures, the pressure drop and the bubble pressure. The pressure drop is the decrease in system pressure across the gas trap. The bubble pressure is the pressure required for air bubbles to pass across the water filled membrane. A significant difference between these pressures is needed to ensure complete capture of air bubbles in a single pass. Bubbles trapped by the device adsorb on the hydrophobic membrane in the interior of the hydrophilic membrane tube. After adsorption, the air is vented due to a pressure drop of approximately 1 atmosphere across the membrane. For water systems, the air is vented to the ambient (cabin). Because water vapor can also transport across the hydrophobic membrane, it is critical that a minimum surface area is used to avoid excessive water loss (would like to have a closed loop for the coolant). The currently used gas traps only provide a difference in pressure drop and bubble pressure of 3-4 psid. This makes the gas traps susceptible to failure at high bubble loading and if gas venting is impaired. One mechanism for the latter is when particles adhere to the hydrophobic membrane, promoting formation of a water layer about it that can blind the membrane for gas transport (Figure 1). This mechanism is the most probable cause for observed failures with the existing design. The objective of this project was to devise a strategy for choosing new membrane materials (database development and procedure), redesign of the gas trap to mitigate blinding effects, and to develop a design that can be used in ammonia and Freon-21 coolant loops.

Ritchie, Stephen M. C.↗

NASA NTRS1998

Pump, and earth-testable spacecraft capillary heat transport loop using augmentation pump and check valves

A spacecraft includes heat-generating payload equipment, and a heat transport system with a cold plate thermally coupled to the equipment and a capillary-wick evaporator, for evaporating coolant liquid to cool the equipment. The coolant vapor is coupled to a condenser and in a loop back to the evaporator. A heated coolant reservoir is coupled to the loop for pressure control. If the wick is not wetted, heat transfer will not begin or continue. A pair of check valves are coupled in the loop, and the heater is cycled for augmentation pumping of coolant to and from the reservoir. This augmentation pumping, in conjunction with the check valves, wets the wick. The wick liquid storage capacity allows the augmentation pump to provide continuous pulsed liquid flow to assure continuous vapor transport and a continuously operating heat transport system. The check valves are of the ball type to assure maximum reliability. However, any type of check valve can be used, including designs which are preloaded in the closed position. The check valve may use any ball or poppet material which resists corrosion. For optimum performance during testing on Earth, the ball or poppet would have neutral buoyancy or be configured in a closed position when the heat transport system is not operating. The ball may be porous to allow passage of coolant vapor.

Baker, David↗

NASA NTRS2010

Validation of Supersonic Film Cooling Modeling for Liquid Rocket Engine Applications

Topics include: upper stage engine key requirements and design drivers; Calspan "stage 1" results, He slot injection into hypersonic flow (air); test articles for shock generator diagram, slot injector details, and instrumentation positions; test conditions; modeling approach; 2-d grid used for film cooling simulations of test article; heat flux profiles from 2-d flat plate simulations (run #4); heat flux profiles from 2-d backward facing step simulations (run #43); isometric sketch of single coolant nozzle, and x-z grid of half-nozzle domain; comparison of 2-d and 3-d simulations of coolant nozzles (run #45); flowfield properties along coolant nozzle centerline (run #45); comparison of 3-d CFD nozzle flow calculations with experimental data; nozzle exit plane reduced to linear profile for use in 2-d film-cooling simulations (run #45); synthetic Schlieren image of coolant injection region (run #45); axial velocity profiles from 2-d film-cooling simulation (run #45); coolant mass fraction profiles from 2-d film-cooling simulation (run #45); heat flux profiles from 2-d film cooling simulations (run #45); heat flux profiles from 2-d film cooling simulations (runs #47, #45, and #47); 3-d grid used for film cooling simulations of test article; heat flux contours from 3-d film-cooling simulation (run #45); and heat flux profiles from 3-d and 2-d film cooling simulations (runs #44, #46, and #47).

Morris, Christopher I.↗

NASA NTRS1947

A Simplified Theory of Porous Wall Cooling

The rate of heat transfer between a fluid stream in turbulent flow and a smooth, solid wall is largely controlled by the relatively high resistance of the laminar sublayer next to the wall. Although this laminar layer ii extremely thin, heat can be transferred through it only by molecular diffusion. Hence the resistance of this layer is very much greater than for a layer the same thickness farther out in the stream where turbulent exchange is the controlling factor. The thickness of the laminar layer is difficult to define precisely, since there is a gradual transition to the turbulent flow outside, but for the usual scale of many engineering applications almost half the temperature difference between the fluid and the wall occurs in a layer of a few thousands of an inch in thickness. When the wall is made of porous material and a coolant gas is forced through the wall into the stream, it has been found that a very small flow rate of the coolant is remarkably effective in keeping the wall at a low temperature. The coolant flow rate required is such as to give an average velocity normal cooling wall of the order of 1 per cent of the main stream velocity. This flow rate is so low that clearly the injected gas must act as an insulator rather than as a normal coolant. Because of its relatively low velocity, the injected gas can have very little influence on heat convection or momentum transfer in the turbulent stream, and its effect must be confined to the laminar sublayer. The possible influence of the coolant flow on the thickness of the laminar layer will be discussed in Section V.

Rannie, W. D.↗

DOE OSTI2020

Acoustic diagnostic system and method for superconducting devices operating in gas or liquid

A novel system and method for detecting a quench of a superconducting conductor and detecting abnormal behavior of a superconducting conductor using acoustic sensor technology in the coolant of a superconducting cable and/or magnet is disclosed. This system and method is not only limited to use for superconductors, but also may be used for any device disposed in liquid and gas. Acoustic sensors are installed along a coolant space of a superconducting conductor and monitor coolant conditions. By monitoring acoustic changes, temperature changes or coolant flow disruption can be detected very quickly by an acoustic sensor array. By disposition of the acoustic sensor array in a coolant flow channel, the acoustic sensor system can quickly detect a local condition, such as the thermal status (temperature) of a superconducting cable and magnet with precise spatial resolution.

75 CONDENSED MATTER PHYSICS, SUPERCONDUCTIVITY AND↗