Yuzhen Lin

Measured Film Cooling Effectiveness of Three Multihole Patterns

As an advanced cooling scheme to meet increasingly stringent combustor cooling requirements, multihole film cooling has received considerable attention. Experimental data of this cooling scheme are limited in the open literature in terms of different hole patterns and blowing ratios. The heat-mass transfer analogy method was employed to measure adiabatic film cooling effectiveness of three multihole patterns. Three hole patterns differed in streamwise row spacing (S), spanwise hole pitch (P), and hole inclination angle (α), with the first pattern S/P=2 and α=30°, the second S/P=1 and a=30°, and the third SIP=2 and a =150°. Measurements were performed at different blow ratios (M=1-4). Streamwise coolant injection offers high cooling protection for downstream rows. Reverse coolant injection provides superior cooling protection for initial rows. The effect of blowing ratio on cooling effectiveness is small for streamwise injection but significant for reversion injection.

Investigation of Film Cooling Effectiveness of Full-Coverage Inclined Multihole Walls With Different Hole Arrangements

An experimental and numerical investigation of adiabatic film cooling effectiveness was conducted on four full-coverage inclined multihole walls with different hole arrangements. The hole geometrical patterns and the test conditions were chosen to be representative of film cooling designs for modern aeroengine combustor liners. The four hole arrangements were grouped into two types based on lateral hole pitch ( P ) and streamwise row spacing ( S ). One type included two test plates which had the same S and P (S/P = 2) and compound angle (β = 0 deg) but different hole inclination angles ( α ) (30 and 150 deg ). The other type included two test plates which had the same S and P (but S/P = 1) and inclination angle (α = 30 deg) but different compound angles (0 deg and 50 deg). Heat-mass transfer analogy method was employed to investigate the adiabatic film cooling effectiveness of these multihole walls with typical blowing ratios for aeroengine combustors. The numerical simulation was performed to characterize the flowfield and temperature distribution, aiming to further understand the film cooling mechanisms. The experimental results indicated that blowing ratio within the range from 1 to 4 had negligible influence on adiabatic film cooling effectiveness (η) in the case of concurrent coolant injection while hole arrangement had large effect on η. But the blowing ratio within the range from 1 to 4 had large effect on the film cooling effectiveness for the counterflow film cooling scheme. The numerical results were compared with experimental data and fairly good agreement was obtained. The numerical simulation revealed the flow structure, particularly exhibiting significant influence of the interaction between mainstream flow and coolant jets on η. With validation by experimental data, film cooling numerical simulation seems quite helpful in selecting optimum multihole arrangement for modern combustor liner design.Copyright

Experimental Investigation on Ignition Performance of LESS Combustor

In the design of next-generation civil aviation gas turbine combustors, there is high demand to improve the efficiency of combustion technology to decrease the amount of fuel consumed and to reduce the emissions in an effort to lessen the environmental impacts. This paper introduces a novel, ultra-low emissions combustor, namely L ow E mission S tirred S wirl (LESS) combustor, employing the lean premixed prevaporized (LPP) approach. The LESS combustor is a single annular layout. Its dome is comprised of two stages — the pilot stage and the main stage. The pilot stage is a typical swirl cup design which uses a pressure swirl atomizer with dual counter-rotating radial swirlers to atomize the fuel and form a diffusion flame, and is located in the centerline of the combustion chamber. The main stage surrounding coaxially the pilot stage includes one annulus as premixer and 14 plain orifice atomizers with 14 small dual counter-rotating radial swirlers arranged uniformly on the dome of the annulus, which lead to the main premixed flame. Five different igniter locations are chosen according to the CFX-simulated non-reacting flow field of a simplified mainstage combustor. Only the pilot pressure swirl atomizer is operated in the present ignition performance tests. The model combustor is a single module rectangular combustor with normal inlet temperature and normal inlet pressure. Under the test conditions of air pressure drop of 0.5%–9%, the ignition performance of the model LESS combustor is analyzed. The lean lightoff fuel/air ratio (LLO FAR), characterizing the ignition performance of a combustor, is investigated herein. In addition, the effects of igniter locations and pilot fuel nozzles on LLO FAR are studied. Specific to the LESS combustor, the igniter location has minor effect on the LLO FAR values. However, as expected, the combustor dome pressure drop and attendant reference velocity along with spray SMD impact LLO FAR. Furthermore, CFX-simulated results of the flow field, spray characteristics, and gas-liquid interactions under the typical condition of combustor operation are presented and discussed to provide insight into the ignition processes and performance.© 2011 ASME

Injection of Supercritical Aviation kerosene fuel into quiescent atmospheric environment

*† ‡ § ** The injection of RP-3 aviation kerosene at conditions near and above the critical point into a quiescent atmospheric environment is experimentally studied, with special emphasis on the effects of fuel injection temperature and pressure on jet characteristics, shock structures, and phase transition. Visualization of the near-field jets and characterization of the shock/jet structures are performed using schlieren imaging. Flow visualization shows fuel condensation phenomenon when supercritical kerosene is injected at reduced temperatures near unity. For the supercritical kerosene injection conditions investigated, ideal gas-like expansion and internal shock structures are observed. The jet expansion angle as well as the axial location and the size of the Mach disk are noted to increase as the injection pressure is increased. Furthermore, while the injection temperature has little effect on the jet structure, it affects the phase transition processes, as demonstrated in the thermodynamic phase diagram calculated using a three-species kerosene surrogate.

Injection of Liquid Kerosene Into a High-Pressure Subsonic Air Crossflow From Normal Temperature to Elevated Temperature

In the present work, injection of liquid kerosene into a high-pressure subsonic air crossflow was investigated experimentally. Tests were conducted at air pressures up to 2.0 MPa and at air temperatures from normal temperature to elevated temperature. Liquid kerosene was injected at room temperature through a 0.5 mm diameter plain orifice. Schlieren imaging technique was used for jet structure visualization, from which the jet penetration trajectory was determined by the image processing. For the conditions tested, a correlation of jet penetration trajectory was developed, with momentum flux ratio, Weber number and crossflow temperature ratio as independent variables. Upper surface trajectories of kerosene spray under different test conditions were compared. Experimental and analytical results showed that the penetration trajectory of liquid kerosene under higher air temperature was greater than that under normal temperature, while momentum ratios were the same.Copyright

Comparison of Flame Dynamics at Stable and Near-LBO Conditions for Swirl-Stabilized Kerosene Spray Combustion

An experimental investigation was conducted to characterize the flame structures and dynamics at stable and near-lean blowout (LBO) conditions. The current experiments were carried out using a laboratory-scale aero-combustor with an internally-staged dome. The internally-staged injector consisted of pilot and main swirlers, and the pilot swirler was fueled with Chinese kerosene RP-3 while the main injector was chocked. The resulting spray flame was confined within a quartz tube under atmosphere pressure. In the present study, the influence of swirl intensity of the pilot swirler was also investiagted. The OH* chemiluminescence of the flame was recorded by a high-speed camera at a frequency of 2000 Hz. From the high-speed OH* images, the reaction zone was marked and the fluctuation of the reaction zone along axial direction was observed, showing that it became stronger at near-LBO condition than at stable condition. Proper Orthogonal Decomposition (POD) analysis was further used to provide insights into the characteristics of flame dynamics. Based on the POD results, the difference of the flame dynamics between the stable and near-LBO combustion was distinct. While the major Mode l of the flame under stable condition was rotation representing the rotation motion in the swirl flame, at near-LBO condition the flame dynamics included three modes — vibration, rotation, and flame shedding. In addition, for swirl-stabilized kerosene spray combustion investigated herein, the fluctuation of the reaction zone in axial direction became stronger with decreasing equivalence ratio when approaching LBO, and the POD analysis indicated that the Mode l of flame dynamics transitions from the rotation mode to the vibration mode. Although the change of pilot swirl number was found to have little influence on the Mode l of flame dynamics, it was noted to vary the fluctuation energy of the flame.Copyright

Influence of the Arrangement of Dilution Holes for Dilution Mixing in a Three-Injector Reverse Flow Combustor

The exit temperature profile has a great effect on the reliability and security in a gas turbine. In this paper, the exit temperature profile of a small engine reverse-flow combustor with three injectors test module was experimentally obtained to qualitatively analyze the influence of the dilution hole distribution. The test model was a three-injector rectangular reverse-flow combustor with swirling flow atomizing. A 1D moving thermocouple rake was used to measure the global exit temperature profile of the combustor. The pressure was at ambient pressure with the inlet temperature was 290K. The FAR was in the range of 0.03. The dilution holes were in opposed and staggered arrangements. The experimental results showed that the exit temperature profile was obviously influenced by the dilution holes. Compared with the opposed dilution jets, the staggered dilution jets provided more uniform circumferential exit temperature profile, but a little higher pattern factor of 0.1725. The numerical results showed that the staggered dilution jets generated a larger scale counter-rotating vortex pairs. The inner and outer jets not only did not interact with each other (especially at the outer ring of combustor), but also filled the intermediate regions of dilution jets, resulting in a higher gas mixing rate. Consequently, the staggered dilution jets provided a better mixing performance for the outer ring of combustor.Copyright

Experimental Study on NOx and CO Emissions of Aviation Kerosene and Coal-to-Liquid Synthetic Aviation Fuel in a Jet Stirred Combustion Reactor

The ever increasing consumption of non-renewable fossil fuels for global economic development leads to serious energy crisis and environmental pollution. Consequently, new alternative fuels and high-efficiency combustion are required to aid the sustainable development of human society. The present paper took the RP-3 aviation kerosene and coal-to-liquid synthetic aviation fuel (manufactured through the Fischer Tropsch process., FT) for object, and experimentally investigated the influences of pressure, inlet temperature and equivalence ratio on the productions of NOx and CO in a jet stirred combustion reactor. The tests were performed under the pressures of 2bar and 3bar, and inlet air temperatures of 550K and 650K, respectively. The equivalence ratio ranged from 0.5 to 1.2. The mean residence time was approximately 8ms. Probe sampling followed by on-line emissions analyzer permitted to measure the concentration of the products. The experimental results show that these two fuels obey the same law with the variations of pressures, inlet temperatures and equivalence ratios. The NOx production increases with the pressure and inlet temperature increasing. The CO decreases with the pressure increasing, while slightly increases with the inlet temperature increasing. Numerical simulations were also performed to investigate the combustion products of these two fuels in the jet stirred combustion reactor. Two PSRs were introduced to simulate the jet flame region and post flame in the recirculation region, respectively. The combustion products of second PSR (PSR2) agreed well with the experimental results by regulating the volume ratio of first PSR (PSR1). Based on the reaction pathway analysis of NO production in present state, it is considered that for these two fuels the NOx production is led by the thermal NO above the equivalence ratio of 0.65, while by the N2O at lower equivalence ratios. With the application of the present alternative fuel and its reaction mechanism, the experimental results of aviation kerosene and Coal-to-Liquid synthetic aviation fuel can be predicted well within a certain state, which requires a further verification in a wider range. Furthermore, the numerical results show that the NO release is insensitive to the reaction components within present experimental states.Copyright

Evaluation of Combustion Performance of a Coal-Derived Synthetic Jet Fuel

For application to aircraft turbines, the present work experimentally examines the physical and combustion-related properties of an F-T synthetic jet fuel relative to the Chinese standard jet fuel, RP-3. This fuel, derived from coal feedstock, is characterized in terms of its physical properties such as density, flash point, freezing point, surface tension, viscosity, and heating value in accordance with Chinese National Standard Testing Methods. Subsequently, several important characteristics relevant to its use in aircraft turbine engines are investigated using a single cup model combustor rig, including atomization, ignition, blowout, and exhaust emissions experiments are carried out. Preliminary results suggest that the use of coal-derived synthetic jet fuel will not result in adverse effects on the performance of an aircraft turbine combustor relative to conventional aviation kerosene. These initial results support the conclusion that full-scale engine testing is warranted to further investigate the performance of F-T synthetic jet fuels in practical systems, and to determine its ability to act as a “drop-in” replacement for traditional aviation fuel.Copyright

Injection of Subcritical and Supercritical Aviation Kerosene Into a High-Temperature and High-Pressure Crossflow

In the present experimental study, injection of subcritical and supercritical kerosene into a high-temperature and high-pressure subsonic crossflow was investigated. Visualization and characterization of the jet structures were performed using schlieren imaging, from which the jet penetration trajectory was determined. For the conditions tested, a correlation of jet penetration trajectory was developed, with momentum ratio as the primary parameter. An analysis based on one-dimensional isentropic flow was also conducted to calculate the flow parameter variations in the nozzle and along the jet trajectory. Using a three-component kerosene surrogate, the phase transition processes for supercritical and subcritical kerosene jets were illustrated in the thermodynamic phase diagram. Experimental and analytical results demonstrated that the behavior and penetration of supercritical kerosene injection into high temperature and high pressure crossflow were closer to those of the case with gas jet injecting into a gas crossflow than the case with liquid kerosene injection.© 2011 ASME