Fuel Influence on Targeted Gas Turbine Combustion Properties: Part I — Detailed Diagnostics

2014 ◽  
Author(s):  
Thomas Mosbach ◽  
Victor Burger ◽  
Barani Gunasekaran
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Leading Edge ◽  
Operating Conditions ◽  
Optical Measurements ◽  
Test Facility ◽  
Research Centre ◽  
High Speed Imaging ◽  
Combustion Properties ◽  
First Results

The threshold combustion performance of different fuel formulations under simulated altitude relight conditions were investigated in the altitude relight test facility located at the Rolls-Royce plc. Strategic Research Centre in Derby, UK. The combustor employed was a twin-sector representation of an RQL gas turbine combustor. Eight fuels including conventional crude-derived Jet A-1 kerosene, synthetic paraffinic kerosenes (SPKs), linear paraffinic solvents, aromatic solvents and pure compounds were tested. The combustor was operated at sub-atmospheric air pressure of 41 kPa and air temperature of 265 K. The temperature of all fuels was regulated to 288 K. The combustor operating conditions corresponded to a low stratospheric flight altitude near 9 kilometres. The experimental work at the Rolls-Royce (RR) test-rig consisted of classical relight envelope ignition and extinction tests, and ancillary optical measurements: Simultaneous high-speed imaging of the OH* chemiluminescence and of the soot luminosity was used to visualize both the transient combustion phenomena and the combustion behaviour of the steady burning flames. Flame luminosity spectra were also simultaneously recorded with a spectrometer to obtain information about the different combustion intermediates and about the thermal soot radiation curve. This paper presents first results from the analysis of the weak extinction measurements. Further detailed test fuel results are the subject of a separate complementary paper [1]. It was found in general that the determined weak extinction parameters were not strongly dependent on the fuels investigated, however at the leading edge of the OH* chemiluminescence intensity development in the pre-flame region fuel-related differences were observed.

Fuel Composition Influence on Gas Turbine Ignition and Combustion Performance

2015 ◽  
Author(s):  
Thomas Mosbach ◽  
Victor Burger ◽  
Barani Gunasekaran
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Jet Fuel ◽  
Synthetic Jet ◽  
Operating Conditions ◽  
Optical Measurements ◽  
Test Facility ◽  
Research Centre ◽  
Combustion Performance ◽  
Insight Into

The influence of different jet fuel compositions on aviation gas turbine combustion performance was investigated. Eight fuels including conventional crude-derived Jet A-1 kerosene, fully synthetic Jet fuel, synthetic paraffinic kerosenes, linear paraffinic solvents, aromatic solvents and pure compounds were tested. The tests were performed in the altitude relight test facility located at the Rolls-Royce Strategic Research Centre in Derby (UK). The combustor employed was a twin-sector representation of an RQL gas turbine combustor. The combustor was operated at sub-atmospheric air pressure of 41 kPa and air temperature of 265 K. The temperature of the fuels was regulated to 288 K. The combustor operating conditions corresponded to a simulated low stratospheric flight altitude near 9,000 metres. The experimental work at the Rolls-Royce (RR) test-rig consisted of classical relight envelope ignition and extinction tests, and ancillary optical measurements: Simultaneous high-speed imaging of the OH* chemiluminescence and of the soot luminescence was applied to obtain spatial and temporal resolved insight into the ongoing processes. Optical emission spectroscopy was also applied simultaneously to obtain spectral and temporal resolved insight into the flame luminescence. First results from the analysis of the OH* chemiluminescence and detailed fuel analysis results were presented in previous papers [1, 2]. This article presents further results from the analysis of the soot luminescence imaging and flame spectra. It was found in general that the combustion performance of all test fuel formulations was comparable to regular Jet A-1 kerosene. Fuel related deviations, if existent, are found to be small.

Characterisation of Waves and Ligaments in Films Close to an Aeroengine Ball Bearing

2019 ◽  
Author(s):  
Jee Loong Hee ◽  
Kathy Simmons ◽  
David Hann ◽  
Michael Walsh
Keyword(s):  
Flow Rate ◽  
High Speed ◽  
Ball Bearing ◽  
Capillary Waves ◽  
Operating Conditions ◽  
Research Centre ◽  
Shaft Speed ◽  
High Speed Imaging ◽  
Liquid Phases ◽  
Mass Transfers

Abstract Surface waves are observed in many situations including natural and engineering applications. Experiments conducted at the Gas Turbine and Transmissions Research Centre (G2TRC) used high speed imaging to observe multiscale wave structures close to an aeroengine ball bearing in a test rig. The dynamic behavior and scale of the waves indicate that these are shear-driven although highly influenced by gravity at low shaft speed. To understand the interactions between gas and liquid phases including momentum and mass transfers, characterization of the observed waves and ligaments was undertaken. Waves were studied at surfaces close to the ball bearing and ligaments were assessed near the cage. Characterization was in terms of frequency and wavelength as functions of speed, flow-rate, bearing axial load and gravity. The assessments confirmed the existence of gravity-capillary waves and capillary waves. Gravity-capillary waves were measured to have a longer mean wavelength on the co-current side of the bearing (gravity and shear acting together) compared to the counter-current side (gravity and shear opposing). Using a published definition of critical wavelength (λcrit), measured wavelengths at 3,000 rpm were 2.56λcrit on the co-current side compared to 1.86λcrit at the countercurrent location. As shaft speed increases, wavelength reduces with transition to capillary waves occurring at around 0.83λcrit. At shaft speeds beyond 5000 rpm, capillary waves were fully visible and the wavelength was obtained as 0.435λcrit. Flow-rate and load did not significantly influence wavelength. Wave frequency was found to be proportional to shaft speed. The gravity-capillary waves had frequencies within the range 13–25 Hz while capillary waves exhibited a frequency well beyond 100 Hz. The frequencies are highly fluctuating with no effect of load and flow rate observed. Ligaments were characterized using Weber number and Stability number. The number of ligaments increased with shaft speed. A correlation for ligament number based on operating conditions is proposed.

Auto-Ignition of In-Line Injected Hydrogen/Nitrogen Fuel Mixtures at Reheat Combustor Operating Conditions

2015 ◽  
Author(s):  
Christoph Schmalhofer ◽  
Peter Griebel ◽  
Michael Stöhr ◽  
Manfred Aigner ◽  
Torsten Wind
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Design Guidelines ◽  
Operating Conditions ◽  
Fuel Injector ◽  
High Speed Imaging ◽  
Capture Process ◽  
Carrier Medium ◽  
Auto Ignition ◽  
Ignition Limits

De-carbonization of the power generation sector becomes increasingly important in order to achieve the European climate targets. Coal or biomass gasification together with a pre-combustion carbon capture process might be a solution resulting in hydrogen-rich gas turbine (GT) fuels. However, the high reactivity of these fuels poses challenges to the operability of lean premixed gas turbine combustion systems because of a higher auto-ignition and flashback risk. Investigation of these phenomena at GT relevant operating conditions is needed to gain knowledge and to derive design guidelines for a safe and reliable operation. The present investigation focusses on the influence of the fuel injector configuration on auto-ignition and kernel development at reheat combustor relevant operating conditions. Auto-ignition of H2-rich fuels was investigated in the optically accessible mixing section of a generic reheat combustor. Two different geometrical in-line configurations were investigated. In the premixed configuration, the fuel mixture (H2 / N2) and the carrier medium nitrogen (N2) were homogeneously premixed before injection, whereas in the co-flow configuration the fuel (H2 / N2) jet was embedded in a carrier medium (N2 or air) co-flow. High-speed imaging was used to detect auto-ignition and to record the temporal and spatial development of auto-ignition kernels in the mixing section. A high temperature sensitivity of the auto-ignition limits were observed for all configurations investigated. The lowest auto-ignition limits are measured for the premixed in-line injection. Significantly higher auto-ignition limits were determined in the co-flow in-line configuration. The analysis of auto-ignition kernels clearly showed the inhibiting influence of fuel dilution for all configurations.

Assessment of the Oil Scoop Capture Efficiency in High Speed Rotors

2018 ◽  
Vol 141 (1) ◽  
Author(s):  
Paloma Paleo Cageao ◽  
Kathy Simmons ◽  
Arun Prabhakar ◽  
Budi Chandra
Keyword(s):  
High Speed ◽  
Geometric Model ◽  
Leading Edge ◽  
Operating Conditions ◽  
Capture Efficiency ◽  
Direct Access ◽  
High Speed Imaging ◽  
Speed Range ◽  
Temperature And Pressure ◽  
Visualization Techniques

Experimental research was conducted into a scooped rotor system that captures oil from a stationary jet and directs it through passages within the shaft to another axial location. Such a system has benefits for delivering oil via under-race feed to aeroengine bearings where direct access is limited. Oil capture efficiency was calculated for three jet configurations, a range of geometric variations relative to a baseline and a range of operating conditions. Flow visualization techniques yielded high-speed imaging in the vicinity of the scoop leading edge. Overall capture efficiency depends on the amount of oil initially captured by the scoop that is retained. Observation shows that when the jet hits the tip of a scoop element, it is sliced and deflected upward in a “plume.” Ligaments and drops formed from this plume are not captured. In addition, some oil initially captured is flung outward as a consequence of centrifugal force. Although in principle capture of the entire supply is possible over most of the shaft speed range, as demonstrated by a simplified geometric model, in practice 60–70% is typical. Significant improvement in capture efficiency was obtained with a lower jet angle (more radial) compared to baseline. Higher capture efficiencies were found where the ratio of jet to scoop tip speed was lower. This research confirms the capability of a scoop system to capture and retain delivered oil. Additional numerical and experimental work is recommended to further optimize the geometry and increase the investigated temperature and pressure ranges.

Investigation of GTL-Like Jet Fuel Composition on GT Engine Altitude Ignition and Combustion Performance: Part II—Detailed Diagnostics

2011 ◽  
Author(s):  
Thomas Mosbach ◽  
Gregor C. Gebel ◽  
Patrick Le Clercq ◽  
Reza Sadr ◽  
Kumaran Kannaiyan ◽  
...  
Keyword(s):  
High Speed ◽  
Operating Conditions ◽  
Research Centre ◽  
Fuel Injector ◽  
Fuel Temperature ◽  
High Speed Imaging ◽  
Combustion Behavior ◽  
Combustion Performance ◽  
Ignition And Combustion ◽  
Spectrally Resolved

The ignition and combustion performance of different synthetic paraffinic kerosenes (SPKs) under simulated altitude relight conditions were investigated at the altitude relight test rig at the Rolls-Royce Strategic Research Centre in Derby. The conditions corresponded to a low stratospheric flight altitude between 25,000 and 30,000 feet. The combustor under test was a twin-sector representation of an advanced gas turbine combustor and fuel injector. Five different SPKs and Jet A-1 were tested at different mass flow rates of air and fuel, and at two different sub-atmospheric air pressures and temperatures. The fuel temperature was kept approximately constant. Simultaneous high-speed imaging of the OH* and CH* chemiluminescence, and of the broadband luminosity was used to visualize both the transient flame initiation phenomena and the combustion behavior of the steady burning flames. In addition, flame luminosity spectra were recorded with a spectrometer to obtain spectrally resolved information concerning the different chemiluminescence bands and the soot luminosity. These investigations were performed in conjunction with the comparative evaluation of the ignition and stability regimes of the five SPKs, which is the subject of a separate complementary paper [1]. We found that the observed flame initiation phenomena, the overall combustion behavior and the different ratios of the chemiluminescence from the OH*, CH* and C2* radicals were not strongly dependent on the fuels investigated. But, the SPK flames showed for all combustor operating conditions significantly lower soot luminosities than the corresponding Jet A-1 flames, indicating a potential benefit of the SPK fuels.

Influence of Carrier Air Preheating on Autoignition of Inline-Injected Hydrogen-Nitrogen Mixtures in Vitiated Air of High Temperature

2017 ◽  
Author(s):  
Christoph A. Schmalhofer ◽  
Peter Griebel ◽  
Manfred Aigner
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Gas Turbines ◽  
High Speed ◽  
Spatial Development ◽  
Design Guidelines ◽  
Operating Conditions ◽  
High Speed Imaging ◽  
Temporal And Spatial ◽  
Fuel Mixtures

Gas turbines will play a significant role in future power generation systems because they provide peak capacity due to their fast start-up capability and high operational flexibility. However, in order to meet the COP 21 goals, de-carbonization of as turbine fuels is required. Compared to natural gas operation, autoignition and flashback risks in gas turbines operated on hydrogen-rich fuels are higher which has to be taken into account for a proper gas turbine design. From investigations of these phenomena at relevant operating conditions with appropriate measurement techniques, e.g. high-speed imaging, the understanding of the non-stationary processes occurring during autoignition can be improved and design guidelines for a safe and reliable gas turbine operation can be derived. The present study investigates the influences of elevated carrier-air preheating temperatures and hydrogen fuel volume fractions on autoignition at hot gas temperatures higher than 1100 K and pressures of 15 bar. An in-line co-flow injector is used to inject the hydrogen-nitrogen fuel mixtures. The formation, temporal and spatial development of autoignition kernels at high-temperature vitiated air conditions, e.g. relevant to reheat combustor operation, are studied. The experiments were conducted in an optically accessible mixing section of a generic reheat combustor. The hydrogen-nitrogen fuel mixtures of up to 70 vol. % hydrogen are injected in-line into the mixing section along with the carrier-air which was preheated to temperatures between 303 K and 703 K. High-speed imaging was used to detect the autoignition kernels and their temporal and spatial development from luminescence signals. Particle Image Velocimetry measurements were conducted to obtain the velocity distribution in the mixing section at autoignition conditions. The influences of vitiated air temperatures and carrier preheating temperatures on autoignition and flame stabilisation limits are shown, alongside the spatial distribution of different types of autoignition kernels, developing at different stages of the autoignition process. The development of autoignition kernels could be linked to the shear layer development derived from global experimental conditions.

Experimental Inverstigations On the Pressure Fluctuations in the Sealing Gap of Dry Gas Seals with Embedded Pressure Sensors

2021 ◽  
Author(s):  
Jingjing Luo ◽  
Dieter Brillert
Keyword(s):  
High Speed ◽  
Gas Flow ◽  
Static Pressure ◽  
Pressure Sensors ◽  
Operating Conditions ◽  
Test Facility ◽  
Experimental Investigations ◽  
Mechanical Seals ◽  
Process Gas ◽  
Sealing Gap

Abstract Dry gas lubricated non-contacting mechanical seals (DGS), most commonly found in centrifugal compressors, prevent the process gas flow into the atmosphere. Especially when high speed is combined with high pressure, DGS is the preferred choice over other sealing alternatives. In order to investigate the flow field in the sealing gap and to facilitate the numerical prediction of the seal performance, a dedicated test facility is developed to carry out the measurement of key parameters in the gas film. Gas in the sealing film varies according to the seal inlet pressure, and the thickness of gas film depends on this fluctuated pressure. In this paper, the test facility, measurement methods and the first results of static pressure measurements in the sealing gap of the DGS obtained in the described test facility are presented. An industry DGS with three-dimensional grooves on the surface of the rotating ring, where experimental investigations take place, is used. The static pressure in the gas film is measured, up to 20 bar and 8,100 rpm, by several high frequency ultraminiature pressure transducers embedded into the stationary ring. The experimental results are discussed and compared with the numerical model programmed in MATLAB, the characteristic and magnitude of which have a good agreement with the numerical simulations. It suggests the feasibility of measuring pressure profiles of the standard industry DGS under pressurized dynamic operating conditions without altering the key components of the seal and thereby affecting the seal performance.

Life Prediction of Directionally Solidified Air Cooled HPT Gas Turbine Blade Used in a Supersonic Aircraft Using FEM

2013 ◽  
Author(s):  
S. Esakki Muthu ◽  
S. Dileep ◽  
S. Saji Kumar ◽  
D. K. Girish
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Operating Conditions ◽  
Test Facility ◽  
Time To Failure ◽  
Generation Process ◽  
Gas Turbine Blade ◽  
Directionally Solidified ◽  
Life Estimation ◽  
Supersonic Aircraft

Life estimation of Directionally Solidified (DS) MARM-247 HPT gas turbine blade used in a turbofan engine of a supersonic aircraft is presented. These blades were drafted into the engine as a replacement for the polycrystal (NIMONIC) blades since a more efficient, reliable and durable material with high strength and temperature resistance was required to further enhance the life of the turbine blade and the efficiency of the power generation process. The supersonic aircraft is having a repeated mission cycle of a fast acceleration from idle, a 1hr cruise at Mach 1.5 and a fast deceleration to idle. The mission cycle which is a repetition of acceleration, cruise and deceleration cycles can produce wide variety of complex loading conditions which can result in HCF, LCF and creep damage of the turbine blade. Empirical equation of the universal slope developed by Manson was used to estimate the damage component due to LCF. The cumulative stresses and strains due to creep as a function of time was determined using Time hardening rule. Creep data for MARM-247 was correlated using LMP to predict the lives to 1% of creep strain at worst possible combination of temperature and stress value. Damage due to creep per mission cycle was determined using Life fraction Rule proposed by Robinson and Taira. The vibration characteristics of the turbine blade were predicted using Modal analysis. Campbell diagram was plotted to ascertain whether any nozzle passing frequency fall within the working range of the blade. Harmonic analysis was carried out to evaluate the magnitude of the alternating stresses resulting from the blade vibrations at resonance during the acceleration and deceleration cycle. HCF life of the turbine blade was assessed using Goodman diagram. The total damage of the turbine blade per mission cycle due to the above loading was assumed as the combination of the individual damage due to fatigue and creep. Time to failure under combined creep and fatigue damage was estimated using linear damage rule. Non linear features of FEA tool ANSYS12.0 was exploited to calculate the stress distribution, creep, plastic and the total strain encountered by the turbine blade as a function of mission cycle time. The loading spectrum associated with the mission cycle which includes the temperature, gas pressure and the speed profiles were obtained from a sophisticated engine ground test facility which was configured to simulate actual engine operating conditions. The proposed method of cyclic life estimation using FEM was validated by performing various component and engine level tests. A good agreement was observed between the calculated and observed blade lives.

The Effect of Heat Transfer Coefficient Increase on Tip Clearance Control in H.P. Compressors in Gas Turbine Engine

2013 ◽  
Author(s):  
Godwin Ita Ekong ◽  
Christopher A. Long ◽  
Peter R. N. Childs
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Time Constant ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Test Facility ◽  
Turbine Engine

Compressor tip clearance for a gas turbine engine application is the radial gap between the stationary compressor casing and the rotating blades. The gap varies significantly during different operating conditions of the engine due to centrifugal forces on the rotor and differential thermal expansions in the discs and casing. The tip clearance in the axial flow compressor of modern commercial civil aero-engines is of significance in terms of both mechanical integrity and performance. In general, the clearance is of critical importance to civil airline operators and their customers alike because as the clearance between the compressor blade tips and the casing increases, the aerodynamic efficiency will decrease and therefore the specific fuel consumption and operating costs will increase. This paper reports on the development of a range of concepts and their evaluation for the reduction and control of tip clearance in H.P. compressors using an enhanced heat transfer coefficient approach. This would lead to improvement in cruise tip clearances. A test facility has been developed for the study at the University of Sussex, incorporating a rotor and an inner shaft scaled down from a Rolls-Royce Trent aero-engine to a ratio of 0.7:1 with a rotational speed of up to 10000 rpm. The idle and maximum take-off conditions in the square cycle correspond to in-cavity rotational Reynolds numbers of 3.1×106 ≤ Reφ ≤ 1.0×107. The project involved modelling of the experimental facilities, to demonstrate proof of concept. The analysis shows that increasing the thermal response of the high pressure compressor (HPC) drum of a gas turbine engine assembly will reduce the drum time constant, thereby reducing the re-slam characteristics of the drum causing a reduction in the cold build clearance (CBC), and hence the reduction in cruise clearance. A further reduction can be achieved by introducing radial inflow into the drum cavity to further increase the disc heat transfer coefficient in the cavity; hence a further reduction in disc drum time constant.

Experimental study of fibre breakup and shot formation in melt blowing nozzle designs

2020 ◽  
pp. 152808372094927 ◽  
Author(s):  
Ignacio Formoso ◽  
Alejandro Rivas ◽  
Gerardo Beltrame ◽  
Gorka S Larraona ◽  
Juan Carlos Ramos ◽  
...  
Keyword(s):  
Experimental Study ◽  
High Speed ◽  
Flux Ratio ◽  
Adhesive Bond ◽  
Operating Conditions ◽  
Nozzle Geometry ◽  
Dimensionless Temperature ◽  
Melt Blowing ◽  
High Speed Imaging ◽  
Adhesive Bonds

The high demand for quality in the manufacture of absorbent hygiene products requires the adhesive bonds between layers to be as uniform as possible. An experimental study was conducted on two industrial multihole melt blowing nozzle designs used for hot-melt adhesive applications for hygiene products, in order to study two defects that influence the quality of the adhesive bond: fibre breakup, resulting in contamination, and the presence of shots, undesirable lumps that end up in the finished product. To this end, the fibre dynamics were captured at the nozzle exit region by using high-speed imaging. From the results it was observed that die drool is the main source of shot formation, while fibre breakup occurs as a result of applying a sufficiently large force in the direction perpendicular to the fibre. In addition, three dimensionless parameters were defined, the first two being the air-polymer flux ratio and the dimensionless temperature ratio, both of which represent the operating conditions, and the remaining one being the force ratio, which represents the nozzle geometry. The effect of these parameters on fibre breakup and shot formation was studied and the results indicate that both the operating conditions and the nozzle geometry were responsible for the onset of the fibre breakup and for the formation of shots. More precisely, both defects turned out to be dominated by the air-polymer flux ratio and the air tilt angle. The results that emerge from this study are useful for the enhancement of industrial melt blowing nozzles.

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