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

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

Combined pyrolysis and radiochemical gas chromatography for studying the thermal degradation of polymers

Pyrolysis gas chromatography and radioactive tracer techniques have been used independently to study the thermal degradation of polymers. In these laboratories the two techniques have been combined to elucidate some of the mechanisms of the thermal degradation of epoxy resins and polyimides. This paper describes the apparatus developed for this work

The X-ray Properties of Five Galactic Supernova Remnants Detected by the Spitzer GLIMPSE Survey

(Abbreviated) We present a study of the X-ray properties of five Galactic supernova remnants (SNRs) -- Kes 17 (G304.6$+$0.1), G311.5$-$0.3, G346.6$-$0.2, CTB 37A (G348.5$+$0.1) and G348.5$-$0.0 -- that were detected in the infrared by Reach et al. (2006) in an analysis of data from the Galactic Legacy Infrared Mid-Plane Survey Extraordinaire (GLIMPSE) that was conducted by the Spitzer Space Telescope. We present and analyze archival ASCA observations of Kes 17, G311.5$-$0.3 and G346.6$-$0.2, archival XMM-Newton observations of Kes 17, CTB 37A and G348.5$-$0.0 and an archival Chandra observation of CTB 37A. All of the SNRs are clearly detected in the X-ray possibly except for G348.5$-$0.0. Our study reveals that the four detected SNRs all feature center-filled X-ray morphologies and that the observed emission from these sources is thermal in all cases. We argue that these SNRs should be classified as mixed-morphology SNRs (MM SNRs): our study strengthens the correlation between MM SNRs and SNRs interacting with molecular clouds and suggests that the origin of mixed-morphology SNRs may be due to the interactions between these SNRs and adjacent clouds. Our ASCA analysis of G311.5$-$0.3 reveals for the first time X-ray emission from this SNR: the X-ray emission is center-filled within the radio and infrared shells and thermal in nature ($kT$ $\sim$ 0.98 keV), thus motivating its classification as an MM SNR. We find considerable spectral variations in the properties associated with the plasmas of the other X-ray-detected SNRs, such as a possible overabundance of magnesium in the plasma of Kes 17. Finally, we also estimate such properties as electron density $n$$_e$, radiative age $t$$_{rad}$ and swept-up mass $M$$_X$ for each of the four X-ray-detected SNRs.Comment: 78 pages, 26 figures, Astronomical Journal, in pres

Description of the warm core turbine facility recently installed at NASA Lewis Research Center

The two net facilities were installed and operated at their design, or rated conditions. The important feature of both of these facilities is that the ratio of turbine inlet temperature to coolant temperature encountered in high temperature engines can be duplicated at moderate turbine inlet temperature. The limits of the facilities with regard to maximum temperature, maximum pressure, maximum mass flow rate, turbine size, and dynamometer torque-speed characteristics are discussed

Effect of cooling-hole geometry on aerodynamic performance of a film-cooled turbine vane tested with cold air in a two-dimensional cascade

The effect of the orientation and cooling-hole size on turbine-vane aerodynamic losses was evaluated. The contribution of individual vane regions to the overall effect was also investigated. Test configurations were based upon a representative configuration having 45 spanwise rows of holes spaced about the entire vane profile. Nominal hole diameters of 0.0254 and 0.0356 cm and nominal hole orientations of 35 deg, 45 deg, and 55 deg from the local vane surface and 0 deg, 45 deg, and 90 deg from the main-stream flow direction were investigated. Flow conditions and aerodynamic losses were determined by vane-exit surveys of total pressure, static pressure, and flow angle

Performance of a high-work low aspect ration turbine tested with a realistic inlet radial temperature profile

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases

Incidence loss for fan turbine rotor blade in two-dimensional cascade

The effect of incidence angle on the aerodynamic performance of a fan turbine rotor blade was investigated experimentally in a two dimensional cascade. The test covered a range of incidence angles from -15 deg to 10 deg and exit ideal critical velocity ratios from 0.75 to 0.95. The principal measurements were blade-surface static pressures and cross-channel survey of exit total pressure, static pressure, and flow angle. Flow adjacent to surfaces was examined using a visualization technique. The results of the investigation include blade-surface velocity distribution and overall kinetic energy loss coefficients for the incidence angles and exit velocity ratios tested. The measured losses are compared with those from a reference core turbine rotor blade and also with two common analytical methods of predicting incidence loss

Cold-air experimental investigation of a turbine with blade trailing edge coolant ejection. 1: Single-stage turbine

Tests were made on a 0.762-meter-tip-diameter research turbine to determine the effect of blade coolant flow on its aerodynamic performance. Both stator and rotor blades had trailing-edge slots for coolant ejection. The turbine was tested over a range of speed and pressure ratio. High primary efficiencies, calculated on the basis of primary air only, were obtained. The efficiency attained was identical to that reported for the turbine from a previous investigation were only slotted stator blades where incorporated in the turbine and tested. And it also compares with results for the turbine with solid blading. Independently varying the rotor coolant flow showed that rotor cooling imposed a severe penalty on turbine efficiency. The thermodynamic efficiency, which accounts for the ideal energies of both blade coolant flows, decreased linearly with rotor coolant at a rate of about 0.7 percent per percent rotor coolant fraction

Design and cold-air test of single-stage uncooled turbine with high work output

A solid version of a 50.8 cm single stage core turbine designed for high temperature was tested in cold air over a range of speed and pressure ratio. Design equivalent specific work was 76.84 J/g at an engine turbine tip speed of 579.1 m/sec. At design speed and pressure ratio, the total efficiency of the turbine was 88.6 percent, which is 0.6 point lower than the design value of 89.2 percent. The corresponding mass flow was 4.0 percent greater than design