Turbine Fuel Ignition and Combustion Facility for Extremely Low Temperature Conditions

2004 ◽  
Author(s):  
G. Pucher ◽  
W. D. Allan
Keyword(s):  
Data Acquisition ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
High Speed ◽  
Air Cooling ◽  
Test Rig ◽  
Ignition System ◽  
Temperature Conditions ◽  
Air Temperatures ◽  
Mode Of Operation

As the temperature of combustion air and fuels are reduced, the ability to achieve ignition within gas turbine engines becomes increasingly difficult. Several factors share responsibility, related largely to the physical characteristics of fuel emerging from nozzles, whereby an increasing fuel viscosity with temperature reduction results in larger average fuel droplets. The ensuing reduced surface area hinders fuel evaporation within an environment where evaporation is already impeded by low partial pressures due to low ambient temperature conditions and/or depending on the mode of operation, due to a high altitude environment. To study the effects of extremely low air and fuel temperatures on gas turbine fuel ignition performance, a dual mode (namely for cold start and altitude relight) test rig has been designed and commissioned. Its main components include a turbo-jet combustion chamber section, fuel system, ignition system, fuel/air cooling systems, and data acquisition/instrumentation. For airflow within the combustion chamber, two alternate sources are used, depending on the mode of operation. As such, this rig allows key parameters related to gas turbine ignition, such as fuel flow, fuel viscosity, ignition characteristics, airflow, and pressure conditions to be monitored and recorded. Highlights of this test rig include a General Electric J-85 combustion chamber section with quartz windows, fuel and air cooling via cryogenic liquids (LN2 for the fuel, LN2 and LOx for air), fuel and air closed loop temperature control, high speed data acquisition, a gas turbine exciter or, as selected, a custom programmable ignition system. Airflow is provided either by twin 11 HP blowers providing up to 0.5 kg/s of airflow to simulate sea level start conditions, or through the entrainment of high velocity air to simulate relight conditions at up to 21000 feet altitude. This rig is capable of achieving minimum inlet air temperatures and fuel temperatures lower than −45°C. A series of commissioning tests was undertaken with the rig in both ground start and altitude (low pressure) configurations. In order to study viscosity effects on ignition performance, two common gas turbine fuels were utilized, namely JP-4 (F-40) and JP-8 (F-34). Ignition fuel flows as well as lean blowout flows for a stock injector design are presented for these fuels across a matrix of fuel and air temperatures. Conclusions are drawn and future developments are described.

An Investigation of Turbine Erosion by Combustor Generated Carbon in a Light Weight Marine Gas Turbine

1978 ◽  
Author(s):  
R. J. Russell ◽  
J. J. Witton
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Operating Conditions ◽  
Fuel Type ◽  
Test Rig ◽  
Erosion Process ◽  
Marine Application ◽  
Mineral Dusts

A study has been made of the turbine erosion problem encountered in a marinized aero gas turbine which arose from the change of fuel type necessitated by the marine application. The work has involved the development of a technique for collecting carbon shed from the combustion chamber under engine operating conditions. Tests using the collector were made with a single combustor test rig and compared to engine experience. Combustion chamber modifications were developed having low solids emissions and their emissions characterized using the collector. The data from the collector show that smaller particles than hitherto collected can produce significant long-term erosion and that reduction on both size and quantity of particles is necessary to reduce erosion to acceptable levels. The data obtained in this study are compared with other published information on the basic erosion process and erosion in gas turbines by natural mineral dusts. The implications of the results to current and future engines are discussed.

Predicting Flashback Limits of a Gas Turbine Model Combustor Based on Velocity and Fuel Concentration for H2-Air-Mixtures

2016 ◽  
Author(s):  
Matthias Utschick ◽  
Daniel Eiringhaus ◽  
Christian Köhler ◽  
Thomas Sattelmayer
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
High Speed ◽  
Air Flow ◽  
Turbulent Flame ◽  
Flame Speed ◽  
Fuel Injector ◽  
Trailing Edge ◽  
Fuel Concentration ◽  
Turbine Model

This study investigates the influence of the fuel injection strategy on safety against flashback in a gas turbine model combustor with premixing of H2-air-mixtures. The flashback propensity is quantified and the flashback mechanism is identified experimentally. The A2EV swirler concept exhibits a hollow, thick walled conical structure with four tangential slots. Four fuel injector geometries were tested. One of them injects the fuel orthogonal to the air flow in the slots (jet-in-crossflow-injector, JICI). Three injector types introduce the fuel almost isokinetic to the air flow at the trailing edge of the swirler slots (trailing edge injector, TEI). Velocity and mixing fields in mixing zone and combustion chamber in isothermal water flow were measured with High-speed-Particle-Image-Velocimetry (PIV) and Highspeed-Laser-Induced-Fluorescence (LIF). The flashback limit was determined under atmospheric pressure for three air mass flows and 673 K preheat temperature for H2-air-mixtures. Flashback mechanism and trajectory of the flame tip during flashback were identified with two stereoscopically oriented intensified high-speed cameras observing the OH* radiation. We notice flashback in the core flow due to Combustion Induced Vortex Breakdown (CIVB) and Turbulent upstream Flame Propagation (TFP) near the wall dependent on the injector type. The Flashback Resistance (FBR) defined as the ratio between a characteristic flow speed and a characteristic flame speed measures the direction of propagation of a turbulent flame in the flow field. Although CIVB cannot be predicted solely based on the FBR, its distribution gives evidence for CIVB-prone states. The fuel should be injected preferably isokinetic to the air flow along the entire trailing edge in oder to reduce the RMS fluctuation of velocity and fuel concentration. The characteristic velocity in the entire cross section of the combustion chamber inlet should be at least twice the characteristic flame speed. The position of the stagnation point should be tuned to be located in the combustion chamber by adjusting the axial momentum. Those measures lead to safe operation with highly reactive fuels at high equivalence ratios.

New Design of Ignition System of Gas Turbine

2014 ◽  
Vol 592-594 ◽  
pp. 1662-1666
Author(s):  
Rahul Singh ◽  
Amber Jain ◽  
Harish Kumar
Keyword(s):  
Thermal Analysis ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Fuel Mixture ◽  
Gas Turbine Engine ◽  
Ignition System ◽  
Turbine Engine ◽  
New Type ◽  
New System

This paper is all about a new type of ignition system for igniting the air-fuel mixture within combustion chamber of a gas turbine engine. In this system there will a separate ignition inside the primary combustion chamber which will be outside the main combustion chamber and responsible for igniting main source of air/fuel mixture inside the combustion chamber. This system is designed to overcome several problems of present ignition system of gas turbine engine and also thermal analysis of this new system has been shown in this paper.

Advanced Seal Test Rig Validation and Operation

2006 ◽  
Author(s):  
Gerrit A. Kool ◽  
Arjen B. Kloosterman ◽  
Edward R. Rademaker ◽  
Bambang I. Soemarwoto ◽  
Fons M. G. Bingen ◽  
...  
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
High Speed ◽  
High Surface ◽  
Weight Ratio ◽  
Labyrinth Seal ◽  
Test Rig ◽  
Turbine Engine ◽  
System A ◽  
Collaborative Program

Advanced seals have been identified as critical in meeting engine goals for specific fuel consumption, thrust-to-weight ratio, emissions, durability, and operating costs. In a direct effort to reduce the parasitic leakage, a high-temperature, high-speed seal test rig with Active Clearance Control (ACC) has been designed, built and validated by the National Aerospace Laboratory (NLR) in the Netherlands within a collaborative program with Sulzer Metco Turbine Components (SMTC) and Pratt & Whitney (P&W). NLR’s new seal test rig is capable to evaluate seals for the next generation gas turbine engines. It will test air seals (i.e., labyrinth, brush, and new seal concepts) in near gas turbine engine environment conditions of high temperature to 815 °C (1500 °F), high pressure to 2400 kPa (335 psid), high surface speeds to 365 m/s (1200 ft/s). Seal flows for typical engine seal clearances between 0.12 mm (0.005 inch) and 0.65 mm (0.025 inch) can be measured without changing test articles but by using the ACC system. A compressed air facility at the German-Dutch Windtunnel, located at the NLR site, delivers the required compressed clean and dry air. This paper describes the design, the instrumentation, the control system and the validation of the test rig. The rig certification was achieved by validating test measurements using a known three knife-edges stepped labyrinth seal. This paper also addresses the NLR’s CFD and engineering tool development to predict the seal performance.

Predicting Flashback Limits of a Gas Turbine Model Combustor Based on Velocity and Fuel Concentration for H2–Air Mixtures

2016 ◽  
Vol 139 (4) ◽  
Author(s):  
Matthias Utschick ◽  
Daniel Eiringhaus ◽  
Christian Köhler ◽  
Thomas Sattelmayer
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
High Speed ◽  
Air Flow ◽  
Turbulent Flame ◽  
Flame Speed ◽  
Fuel Injector ◽  
Trailing Edge ◽  
Fuel Concentration ◽  
Turbine Model

This study investigates the influence of the fuel injection strategy on safety against flashback in a gas turbine model combustor with premixing of H2–air mixtures. The flashback propensity is quantified and the flashback mechanism is identified experimentally. The A2EV swirler concept exhibits a hollow, thick-walled conical structure with four tangential slots. Four fuel injector geometries were tested. One of them injects the fuel orthogonal to the air flow in the slots (jet-in-crossflow injector (JICI)). Three injector types introduce the fuel almost isokinetic to the air flow at the trailing edge of the swirler slots (trailing edge injector (TEI)). Velocity and mixing fields in mixing zone and combustion chamber in isothermal water flow were measured with high-speed particle image velocimetry (PIV) and high-speed laser-induced fluorescence (LIF). The flashback limit was determined under atmospheric pressure for three air mass flows and 673 K preheat temperature for H2–air mixtures. Flashback mechanism and trajectory of the flame tip during flashback were identified with two stereoscopically oriented intensified high-speed cameras observing the OH* radiation. We notice flashback in the core flow due to combustion-induced vortex breakdown (CIVB) and turbulent flame propagation (TFP) near the wall dependent on the injector type. The flashback resistance (FBR) defined as the ratio between a characteristic flow speed and a characteristic flame speed measures the direction of propagation of a turbulent flame in the flow field. Although CIVB cannot be predicted solely based on the FBR, its distribution gives evidence for CIVB-prone states. The fuel should be injected preferably isokinetic to the air flow along the entire trailing edge in order to reduce the RMS fluctuation of velocity and fuel concentration. The characteristic velocity in the entire cross section of the combustion chamber inlet should be at least twice the characteristic flame speed. The position of the stagnation point should be tuned to be located in the combustion chamber by adjusting the axial momentum. Those measures lead to safe operation with highly reactive fuels at high equivalence ratios.

Influence of Autoignition Kernel Development on the Flame Stabilisation of Hydrogen-Nitrogen Mixtures in Vitiated Air of High Temperature

2018 ◽  
Author(s):  
Christoph A. Schmalhofer ◽  
Peter Griebel ◽  
Manfred Aigner
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Vertical Direction ◽  
Operating Conditions ◽  
Data Set ◽  
Luminescence Signal ◽  
Air Temperatures ◽  
Temperature Regimes ◽  
Depth Analysis ◽  
Nitrogen Mixture

Autoignition and flame stabilisation in a gas turbine combustor presents severe challenges for safe and reliable gas turbine operation as soon as they occur in parts of the combustor that are not designed to sustain higher thermal loads. Especially when operating on highly-reactive fuels like hydrogen, higher autoignition and flashback risk associated with these fuels have to be taken into account. In the present study, flame stabilisation initiated by autoignition events is investigated in an optically accessible mixing duct of a generic reheat combustor at typical reheat conditions. The experiments were conducted at pressures of 15 bar, vitiated air temperatures higher than 1100 K and bulk velocity of 200 m/s. The fuel was a hydrogen-nitrogen mixture with up to 70 vol. % hydrogen and was injected by a coflow inline injector along with preheated carrier air of temperatures up to 623 K. The autoignition-driven flame stabilisation process was investigated by recording the luminescence signal with high-speed cameras and by tracking the temporal and spatial development of autoignition kernels in the mixing duct. A detailed and comprehensive data set could be generated providing the basis for an in-depth analysis of the stabilisation process on time scales down to 0.3 milliseconds, which is fast enough to disclose the small timescales at which the autoignition kernels develop in the mixing section. A stabilising sequence was found to lead to the stabilised flames due to a non-interrupted sequence of autoignition kernels. The stabilising sequence was found behave differently in two different temperature regimes where sequence durations and amounts of kernels differed significantly from each other. A state in which the cross section of the mixing section is fully blocked by one or more kernels in vertical direction could be identified for all operating conditions and the development of subsequent autoignition kernels after the section blockage changed clearly once this state was reached.

Experimental Investigation of the Combustion Characteristics of a Double-Staged FLOX®-Based Combustor on an Atmospheric and a Micro Gas Turbine Test Rig

2015 ◽  
Author(s):  
Jan Zanger ◽  
Thomas Monz ◽  
Manfred Aigner
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Wide Field ◽  
Exhaust Gas ◽  
Combustion System ◽  
Test Rig ◽  
Atmospheric Measurements ◽  
Gas Emissions ◽  
Micro Gas Turbine ◽  
Exhaust Gas Emissions

To establish micro gas turbine (MGT) systems in a wide field of CHP applications, innovative combustion concepts are needed to meet the demands for low exhaust gas emissions, high efficiency and reliability as well as high fuel flexibility. A promising technology for future MGT combustion is the FLOX® concept. The goal of the presented work is to prove the feasibility of a double–staged, FLOX®–based MGT combustion system on a MGT test rig. The paper reports a reliable operating behavior of a Turbec T100 MGT in combination with the new FLOX®–based combustion chamber utilizing natural gas. The measured exhaust gas emissions are compared for different configurations of the combustion chamber and the standard Turbec system. It is shown that the carbon monoxide emissions are reduced whereas the nitrogen oxide emissions exceed the emission levels of the standard MGT burner. However, they still fall far below the German legal limits. For helping to interpret the results of the MGT combustion system, the double–staged combustor is compared to a single–staged FLOX®burner on basis of atmospheric measurements. Here, it is shown that the margin to lean blow–off is substantially increased by the fuel staging. Moreover, it is demonstrated that the exhaust gas emissions of the double–staged combustor could be kept at a similar very low level by applying the staging. Additionally, the overall reaction regions are reported by OH* chemiluminescence imaging as a function of burner air number. Based on this atmospheric study the transfer to MGT conditions is made and appropriate measures are derived to optimize the exhaust gas emissions of the MGT FLOX® combustion system.

Applications of Gas Turbine Plants With Cooled Compressor Intake Air

2001 ◽  
Author(s):  
E. Kakaras ◽  
A. Doukelis ◽  
J. Scharfe
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Cooling System ◽  
Ambient Air ◽  
Climatic Conditions ◽  
Air Cooling ◽  
Iso Standard ◽  
Air Temperatures ◽  
Air Cooling System ◽  
Alternative Scenarios

The operation of gas turbines at ambient air temperatures higher than the ISO standard conditions (15°C) causes performance penalties both in the generated power and the efficiency of the engine. At high inlet-air temperatures, there can be a power loss of more than 20% combined with a significant increase in specific fuel consumption, compared to the ISO standard conditions. Thus, over a long period of time, gas turbines have a lower power output and efficiency than the equipment could actually perform. It is the purpose of this work to present the possibilities and advantages from the integration of an innovative air-cooling system for reducing the gas turbine intake-air temperature. The advantages of this system are demonstrated by examining alternative scenarios of usage, representative of different countries and different climatic conditions.

A Gas Turbine Engine Backup Bearing Operating Beyond 2.5 Million DN

2000 ◽  
Author(s):  
Erik Swanson ◽  
James F. Walton ◽  
Hooshang Heshmat
Keyword(s):  
Gas Turbine ◽  
High Speed ◽  
Gas Turbine Engine ◽  
Magnetic Bearing ◽  
Active Magnetic Bearing ◽  
Test Rig ◽  
Turbine Engine ◽  
Post Test ◽  
Rotor Support ◽  
Auxiliary Bearing

Gas turbine engines and high speed rotating machinery using magnetic bearings require auxiliary and backup bearings for reliability and safety of operation. A 140 mm diameter Zero Clearance Auxiliary Bearing (ZCAB) capable of supporting radial and/or thrust loads of up to 4500 N was designed for an advanced gas turbine engine. The ZCAB was fabricated and tested successfully up to the expected maximum operating speed of 18,000 rpm in a specially configured test rig. The test rig included a 36,000 rpm capable drive motor, a 64 kg rotor which simulates a gas turbine engine shaft dynamics, a damped ball bearing at the drive end and an active magnetic bearing next to the ZCAB. Operation in excess of 240 minutes and 20 transient engagements simulating magnetic bearing failures were completed in the initial tests. Post test inspection revealed minimal wear to the shaft and the ZCAB rollers, whereupon the ZCAB was reassembled for shipment. These preliminary tests confirm the operation and durability of the ZCAB in maintaining rotor support and continued operation even if the primary magnetic bearing support is overloaded or encounters a failure.

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