• Transactions of the Korean Society of Mechanical Engineers
• /
• v.12 no.3
• /
• pp.582-589
• /
• 1988

A rapid expansion of moist air or steam in a supersonic nozzle gives rise to condensation, and the total pressure of the flow is decreased due to this irreversibility of condensation phenomenon. In the present paper, the loss in total pressure during the condensation process has been studied, by numerical analysis and pressure measurement, in the case of moist air expanding in a supersonic nozzle. The effects of the degree of supersaturation at the stagnation condition and expansion rate of the nozzle on the total pressure loss have been studied. The length of the region where the total pressure decreases during the condensation process is longer than that of the nonequilibrium condensation region, and of difference between the length of these two increases with the increase of the degree of supersaturation at the stagnation condition. Furthermore, the larger the expansion rate of the nozzle and the higher the temperature and the degree of supersaturation at the reservoir are, the larger the total pressure loss of the flow becomes. And, it is turned out that the total pressure loss be about 2 to 8 percent in the present study.

• Transactions of the Korean Society of Mechanical Engineers B
• /
• v.29 no.8 s.239
• /
• pp.956-962
• /
• 2005

A three-dimensional computation was conducted to make a study about effects of the inlet boundary layer thickness on the total pressure loss in a low-speed axial compressor operating at the design condition ($\phi=85\%$) and near stall condition($\phi=65\%$). Differences of the tip leakage flow and hub corner-stall induced by the inlet boundary layer thickness enable the loss distribution of total pressure along the span to be altered. At design condition, total pressure losses for two different inlet boundary layers are almost alike in the core flow region but the larger loss is generated at both hub and tip when the inlet boundary layer is thin. At the near stall condition, however, total pressure loss fer the thick inlet boundary layer is found to be greater than that for the thin inlet boundary layer on most of the span except the region near hub and casing. Total pressure loss is scrutinized through three major loss categories in a subsonic axial compressor such as profile loss, tip leakage loss and endwall loss using Denton's loss model, and effects of the inlet boundary layer thickness on the loss structure are analyzed in detail.

• Journal of the Korean Society of Fisheries and Ocean Technology
• /
• v.28 no.2
• /
• pp.157-163
• /
• 1992

In this paper, the relationship between static pressure recovery and turbulent energy was presented in case of swirling flows into a conical diffuser. The distributions of turbulent energy in a diffuser sectional area were measured by a hot wire anemometer. The following conclusion can be drawn from the experiment. Diffuser loss is constituted by a dynamic pressure loss and total pressure loss. The static pressure recovery depends strongly on the total pressure loss. The static pressure recovery depends strongly on the total pressure loss, and the turbulent energy varies inversely as the static pressure recovery coefficient.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.295-298
• /
• 2009

Bent ducts add loss and decrease efficiency. Many researchers have been conducted the performances of bent ducts, but their shapes of inlet and outlet are same. However, in this investigation, the focus is on a bent duct which is annular at the inlet and circular at the outlet. The bent duct of these complex shapes has not been investigated, but has been used in many fields. The performance of such bent duct is investigated under inlet speed 54 m/s and Re = 238,000. Wall static pressure tappings are located surface of the bent duct to measure the static pressure and a probe is traversed at the inlet and outlet of the bent duct to measure the total pressure. As a result, it presents static pressure distribution on the bent duct surface, streamwise velocity profile at inlet and outlet of the bent duct and total pressure loss profile at outlet. In this investigation, the total pressure loss coefficient is 0.243.

• Transactions of the Korean Society of Mechanical Engineers B
• /
• v.20 no.4
• /
• pp.1363-1375
• /
• 1996

Oil-film flow visualizations and three-dimensional flow measurements using a five-hole probe have been conducted to investigate three-dimensional flow characteristics and total pressure losses of a row of film-cooling jets injected in spanwise direction. For several span-to-diameter ratios, experiments are performed in the case of three velocity ratios of 0.5, 1.0 and 1.5. The flow measurements show that downstream flow due to the injection is characterized by a single streamwise vortex instead of a pair of counter-rotating vortices, which appear in the case of streamwise injection, and the vortex strength strongly depends on the velocity ratio. Regardless of the velocity*y ratio, presence of the spanwise film-cooling jets always produces total pressure loss, which is pronounced when the velocity ratio is large. It has also been found that the production of the total pressure loss is closely related to the secondary vortical flow. In addition, effects of the span-to-diameter ratio on the flow and total pressure loss are discussed in detail.

• The KSFM Journal of Fluid Machinery
• /
• v.2 no.1 s.2
• /
• pp.73-80
• /
• 1999

Leakage vortices formed new blade tip causes an increase of total pressure loss near the casing endwall region and as a result, the efficiency of rotor decreases. The reduction of rotor efficiency is related to the size of the tip clearance. In this study, the three-dimensional flowfields in an axial flow rotor were calculated by varying the tip clearance under various flow rates, and the numerical results were compared with experimental ones. The effects of tip clearance and attack angle on the leakage vortex and overall performance, and the loss distributions were investigated through numerical calculations. In this study, tip leakage flow rate and total pressure loss by tip clearance were evaluated using numerical results and approximate equations were presented to evaluate the reduction of rotor efficiency by tip leakage flow.

• The KSFM Journal of Fluid Machinery
• /
• v.2 no.4 s.5
• /
• pp.48-56
• /
• 1999

For the study on loss coefficients of turbine cascade with variation of incidence angle, the wind-tunnel tests were performed under the ranges in velocity of 10 m/s, 15 m/s, 20 m/s and incidence angles from $-20^{\circ}\;to\;20^{\circ}$ by intervals of $5^{\circ}$. Comparing our results with Soderberg's prediction, differences in loss coefficient were $2.5\%\;and\;2.8\%$ each for 10 m/s and 15 m/s. A large disagreement of $30.3\%$ was showed at 20 m/s freestream velocity. The comparisons of these test results with Ainley's prediction showed an $8\%$ difference in the case of 20 m/s freestream velocity. Test results were approximately comparable with Ainley's loss prediction's in incidence angles. Generally, averaged total pressure loss seemed to be decreased as Reynolds number increased. The total pressure loss coefficients were increased parabolically, as incidence angles were increased negatively and positively from $0^{\circ}$, in all speed ranges. At the far low freestream velocities, minimum loss accurred between $-5^{\circ}\;and\;+5^{\circ}$. But this minimum range narrowed the location of this range by shifting to the direction of the angle as freestream velocity was increased.

• Journal of Mechanical Science and Technology
• /
• v.14 no.4
• /
• pp.456-465
• /
• 2000

An experiment was conducted to investigate the aerodynamic losses of high pressure steam turbine nozzle (526A) subjected to a large range of incident angles ($-34^{\circ}\;to\;26^{\circ}$) and exit Mach numbers (0.6 and 1.15). Measurements included downstream Pitot probe traverses, upstream total pressure, and end wall static pressures. Flow visualization techniques such as shadowgraph and color oil flow visualization were performed to complement the measured data. When the exit Mach number for nozzles increased from 0.9 to 1.1 the total pressure loss coefficient increased by a factor of 7 as compared to the total pressure losses measured at subsonic conditions ($M_2<0.9$). For the range of incidence tested, the effect of flow incidence on the total pressure losses is less pronounced. Based on the shadowgraphs taken during the experiment, it' s believed that the large increase in losses at transonic conditions is due to strong shock/ boundary layer interaction that may lead to flow separation on the blade suction surface.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.11a
• /
• pp.145-148
• /
• 2008

Korea Aeronautics Research Institute (KARI) is developing a turbo pump that has 1-stage impulse turbine and relatively high tip clearance for safety. The objective of this research is to investigate the effect of reaction on tip clearance loss in axial turbines. Both cascades were tested in a subsonic wind tunnel. In each cascade, total pressure was measured for tip clearance ranging from 1% to 20% of chord. In results, increasing tip clearance, total pressure loss in reaction turbines is continually increased but impulse turbines keep almost same level of mass averaged total pressure loss. When tip clearance becomes more than 10% of chord, mass-averaged total pressure loss in impulse turbines is less than in reaction. This means that when tip clearance is more than 10% of chord, impulse turbines have better efficiency than reaction turbines.

• 유체기계공업학회:학술대회논문집
• /
• 2004.12a
• /
• pp.419-426
• /
• 2004

A three-dimensional computation was conducted to understand effects of the inlet boundary layer thickness on the loss mechanism in a low-speed axial compressor operating at the design condition(${\phi}=85\%$) and near stall condition(${\phi}=65\%$). At the design condition, the flow phenomena such as the tip leakage flow and hub comer stall are similar independent of the inlet boundary layer thickness. However, when the axial compressor is operating at the near stall condition, the large separation on the suction surface near the casing is induced by the tip leakage flow and the boundary layer on the blade for thin inlet boundary layer but the hub corner stall is enlarged for thick inlet boundary layer. These differences of internal flows induced by change of the boundary layer thickness on the casing and hub enable loss distributions of total pressure to be altered. When the axial compressor has thin inlet boundary layer, the total pressure loss is increased at regions near both casing and tip but decreased in the core flow region. In order to analyze effects of inlet boundary layer thickness on total loss in detail, using Denton's loss models, total loss is scrutinized through three major loss categories in a subsonic axial compressor such as profile loss, tip leakage loss and endwall loss.