Experimental Investigations on Temperature Distribution Along the Liner Wall and Centerline of Small Capacity Gas Turbine Combustion Chamber for Fuel Rich to Fuel Lean Air/Fuel Ratios at Equivalence Ratio of 1

2006 ◽  
Author(s):  
Digvijay B. Kulshreshtha ◽  
S. A. Channiwala
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Fuel Mixture ◽  
Combustion Efficiency ◽  
Experimental Investigations ◽  
Temperature Profiles ◽  
Fuel Ratio ◽  
Wide Range ◽  
Exit Temperature ◽  
Unburned Fuel

The development of the combustion chamber for 20kW gas turbine unit using kerosene type fuel has been undertaken keeping in view the basic requirements of a good combustion chamber, namely, high combustion efficiency, low pressure loss, smooth ignition, wide stability limits, size and shape compatible with engine envelop, low emissions of smoke, unburned fuel and gaseous pollutant species, durability and ease of maintenance. A sophisticated experimental test rig has then been developed to investigate over a wide range of air/fuel ratios for the temperature profiles at the few axial and liner wall locations of this combustion chamber. The range of overall air/fuel ratios considered varies from 22.7396 to 152.4 i.e. Rich Air/Fuel Mixture to Lean Air/Fuel Mixture Range. The temperature profiles for centerline and liner wall for eight different air/fuel ratios are summarized here. The two air/fuel ratios selected are near the designed value of 118.34. It could be concluded from the results that the air/fuel ratio of 122.106 gives the best results for centerline temperature and the liner wall temperature as well as the exit temperature profile. This is very near to the designed air/fuel ratio of 118.34. The temperatures of near 1400 °C achieved at the centerline of the combustion chamber and the liner wall temperatures in the range of 500 °C for lower air/fuel ratio and 300 °C for higher air/fuel ratio certainly ensures safe and reliable operation of this combustion chamber.

Experimental Investigations of Pressure Drop in the Combustion Chamber of Gas Turbine

1989 ◽  
Author(s):  
Marek Dzida ◽  
Krzysztof Kosowski
Keyword(s):  
Pressure Drop ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Operating Conditions ◽  
Experimental Results ◽  
Experimental Investigations ◽  
Relative Pressure ◽  
Wide Range ◽  
Few Data

In bibliography we can find many methods of determining pressure drop in the combustion chambers of gas turbines, but there is only very few data of experimental results. This article presents the experimental investigations of pressure drop in the combustion chamber over a wide range of part-load performances (from minimal power up to take-off power). Our research was carried out on an aircraft gas turbine of small output. The experimental results have proved that relative pressure drop changes with respect to fuel flow over the whole range of operating conditions. The results were then compared with theoretical methods.

Development of Dry Low-NOx Combustor for 300 kW Class Gas Turbine Applied to Co-Generation Systems

2001 ◽  
Author(s):  
Yoichiro Ohkubo ◽  
Osamu Azegami ◽  
Hiroshi Sato ◽  
Yoshinori Idota ◽  
Shinichiro Higuchi
Keyword(s):  
Gas Turbine ◽  
Output Power ◽  
Combustion Efficiency ◽  
Gas Generator ◽  
Turbine Engine ◽  
Partial Load ◽  
Wide Range ◽  
Compressor Inlet ◽  
Endurance Testing ◽  
Turbine Speed

A 300 kWe class gas turbine which has a two-shaft and simple-cycle has been developed to apply to co-generation systems. The gas turbine engine is operated in the range of about 30% partial load to 100% load. The gas turbine combustor requires a wide range of stable operations and low NOx characteristics. A double staged lean premixed combustor, which has a primary combustion duct made of Si3N4 ceramics, was developed to meet NOx regulations of less than 80 ppm (corrected at 0% oxygen). The gas turbine with the combustor has demonstrated superior low-emission performance of around 40 ppm (corrected at 0% oxygen) of NOx, and more than 99.5% of combustion efficiency between 30% and 100% of engine load. Endurance testing has demonstrated stable high combustion performance over 3,000 hours in spite of a wide compressor inlet air temperature (CIT) range of 5 to 35 degree C.. While increasing the gas generator turbine speed, the flow rate of primary fuel was controlled to hold a constant equivalence ratio of around 0.5 in the CIT range of more than 15 C. The output power was also decreased while increasing the CIT, in order to keep a constant temperature at the turbine inlet. The NOx decreases in the CIT range of more than 15 C. On the other hand, the NOx increases in the CIT range of less than 15 C when the output power was kept a constant maximum power. As a result, NOx emission has a peak value of about 40 ppm at 15 C.

Development and Testing of a FuelFlex Dry-Low-NOx Micromix Combustor for Industrial Gas Turbine Applications With Variable Hydrogen Methane Mixtures

2019 ◽  
Author(s):  
H. H.-W. Funke ◽  
N. Beckmann ◽  
S. Abanteriba
Keyword(s):  
Gas Turbine ◽  
Experimental Studies ◽  
Fuel Mixture ◽  
Combustion Efficiency ◽  
Pollutant Emission ◽  
Pure Hydrogen ◽  
Negative Effects ◽  
Low Emission ◽  
Combustion Properties ◽  
Fuel Mixtures

Abstract The negative effects on the earth’s climate make the reduction of the potent greenhouse gases carbon-dioxide (CO2) and nitrogen oxides (NOx) an imperative of the combustion research. Hydrogen based gas turbine systems are in the focus of the energy producing industry, due to their potential to eliminate CO2 emissions completely as combustion product, if the fuel is produced from renewable and sustainable energy sources. Due to the difference in the physical properties of hydrogen-rich fuel mixtures compared to common gas turbine fuels, well established combustion systems cannot be directly applied for Dry Low NOx (DLN) hydrogen combustion. The paper presents initial test data of a recently designed low emission Micromix combustor adapted to flexible fuel operation with variable fuel mixtures of hydrogen and methane. Based on previous studies, targeting low emission combustion of pure hydrogen and dual fuel operation with hydrogen and syngas (H2/CO 90/10 vol.%), a FuelFlex Micromix combustor for variable hydrogen methane mixtures has been developed. For facilitating the experimental low pressure testing the combustion chamber test rig is adapted for flexible fuel operation. A computer-controlled gas mixing facility is designed and installed to continuously provide accurate and homogeneous hydrogen methane fuel mixtures to the combustor. An evaluation of all major error sources has been conducted. In the presented experimental studies, the integration-optimized FuelFlex Micromix combustor geometry is tested at atmospheric pressure with hydrogen methane fuel mixtures ranging from 57 vol.% to 100 vol.% hydrogen in the fuel. For evaluating the combustion characteristics, the results of experimental exhaust gas analyses are applied. Despite the design compromise, that takes into account the significantly different fuel and combustion properties of the applied fuels, the initial results confirm promising operating behaviour, combustion efficiency and pollutant emission levels for flexible fuel operation. The investigated combustor module exceeds 99.4% combustion efficiency for hydrogen contents of 80–100% in the fuel mixture and shows NOx emissions less than 4 ppm corrected to 15 vol.% O2 at the design point.

Assessment of the Operational Performance of Fischer-Tropsch Synthetic-Paraffinic Kerosene in a T63 Gas Turbine Compared to Conventional Jet A-1 Fuel

2009 ◽  
Author(s):  
Nigel Bester ◽  
Andy Yates
Keyword(s):  
Gas Turbine ◽  
Thermal Loading ◽  
Flame Temperature ◽  
Combustion Products ◽  
Combustion Efficiency ◽  
Efficiency Gain ◽  
Engine Efficiency ◽  
Exit Temperature ◽  
The Impact ◽  
Combustor Liner

The performance implications of operating on Synthetic-Paraffinic Kerosene (SPK) were investigated using a RR-Allison T63-A-700 Model 250-C18 B gas turbine and compared to conventional Jet A-1. The SPK was aromatic–free and possessed a greater hydrogen/carbon ratio than petroleum derived Jet A-1. The variation in aromatic content had several implications with respect to soot and NOx emissions. Reduced aromatics also implied a reduction in the radiative heat transfer to the combustor liner. A simple model was used to explore the effect of H/C ratio on the adiabatic flame temperature, the combustor exit temperature and the engine efficiency via the impact on the gas properties and these were compared to the experimental data. It was found that operation with SPK changed directionally toward improving energy extraction via a turbine and an overall efficiency gain of about 1.2% was attained with operation on SPK through increased combustion efficiency, a reduction in liner pressure loss and an improvement in the combustion products properties. A modified combustion liner was fitted to enable the thermal loading on the combustor liner to be investigated and the expected trend with the SPK fuel was confirmed and quantified.

Numerical Simulation as Design Optimization Tool for Gas Turbine Combustion Chambers

2010 ◽  
Author(s):  
Digvijay B. Kulshreshtha ◽  
S. A. Channiwala ◽  
Saurabh B. Dikshit
Keyword(s):  
Numerical Simulation ◽  
Temperature Distribution ◽  
Combustion Chamber ◽  
Design Optimization ◽  
Gas Turbine ◽  
Flow Patterns ◽  
Intermediate Zone ◽  
Experimental Investigations ◽  
Combustion Chambers ◽  
Primary Zone

In present study an attempt has been made through CFD approach using CFX 11 to analyze the flow patterns within the combustion liner and through different air admission holes, namely, primary zone, intermediate zone, dilution zone and wall cooling, and from these the temperature distribution in the liner and at walls as well as the temperature quality at the exit of the combustion chamber are predicted. The design optimization is carried out using the CFD results with validation using experimental investigations.

Effects of Inlet Air Distortion on Gas Turbine Combustion Chamber Exit Temperature Profiles

2015 ◽  
Author(s):  
O. Maqsood ◽  
M. LaViolette ◽  
R. Woodason
Keyword(s):  
Combustion Chamber ◽  
Operating Conditions ◽  
Temperature Fields ◽  
Temperature Profiles ◽  
Toroidal Vortex ◽  
Uniform Temperature ◽  
Sauter Mean Diameter ◽  
Test Rig ◽  
Turbine Performance ◽  
Exit Temperature

Localized damage to turbine inlet nozzles is typically caused by non-uniform temperature distributions at the combustion chamber exit. This damage results in decreased turbine performance and can lead to expensive repair or replacement. A test rig was designed and constructed for the Rolls-Royce Allison 250-C20B dual-entry combustion chamber to investigate the effects of inlet air distortion on the combustion chamber’s exit temperature fields. The rig includes a purposely built water cooled thermocouple rake to sweep the exit plane of the combustion chamber. Test rig operating conditions simulated normal engine cruise conditions by matching the quasi-non-dimensional Mach number, equivalence ratio and Sauter mean diameter. The combustion chamber was tested with an even distribution of inlet air and a 4% difference in airflow at either combustion chamber inlet. An even distribution of inlet air to the combustion chamber did not produce a uniform temperature profile and varying the inlet distribution of air exacerbated the profile’s non-uniformity. The design of the combustion chamber promoted the formation of an oval-shaped toroidal vortex inside the combustion liner, causing localized hot and cool sections separated by 90° that were apparent in the exhaust. Uneven inlet air distributions skewed the oval vortex, increasing the temperature of the hot section nearest the side with the most airflow and decreasing the temperature of the hot section on the opposite side.

scholarly journals Analysis of Efficiency upon Enhancing the Fuel Combustion Completeness in the GTU Burners Using Fuel Gas Heating up

2021 ◽  
pp. 88-98
Author(s):  
Ahiley Pekov ◽  
◽  
Nikolai Bachev ◽  
Alena Shilova ◽  
Oleg Matyunin ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Combustion Efficiency ◽  
Fuel Combustion ◽  
Fuel Gas ◽  
Temperature Changes ◽  
Gas Heating ◽  
Different Temperatures ◽  
Design Characteristics ◽  
Heating Up

One main characteristic of the gas turbine unit (GTU) burner is its fuel combustion completeness, which affects directly the efficiency of the power plant along with CO and unburnt hydrocarbons CnHm emissions. The aim of this work was the research on the application of the fuel heating-up as an alternative method for increasing the fuel combustion completeness and controlling the emission of harmful agents. This goal is achieved by obtaining experimental data on the emissions of CO and NOx at different temperatures of the fuel gas supply to the combustion chamber. The most significant result of the work is the experimentally confirmed possibility of increasing the combustion efficiency (decreasing CO) by heating the fuel gas while maintaining constant gas-dynamic characteristics of the chamber. The significance of the results obtained consists in the experimental confirmation of the combustion quality control only by heating the fuel gas without changing the operating and design characteristics of the combustion chamber. The fuel combustion low completeness can cause the burner unstable operation in the form of the unsteady pre-blowout burning combined with the pressure oscillations in the burner. At present, methods for ensuring the increase in stability and completeness of the fuel combustion are related to the air rate and temperature changes at the inlet. However, the use of these methods can be unwanted because of their causing the decrease in the coefficient of efficiency and in the resource of the ‘hot part’ of the gas-turbine facility.

THERMO ACOUSTIC PROCESSES IN LOW EMISSION COMBUSTION CHAMBER OF GAS TURBINE ENGINE CAPACITY 25 MW

2019 ◽  
pp. 86-90
Author(s):  
Sergey Serbin
Keyword(s):  
Mathematical Models ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Three Dimensional ◽  
Fuel Mixture ◽  
Gas Turbine Engine ◽  
Operational Parameters ◽  
Combustion Chambers ◽  
Turbine Engine ◽  
Low Emission

The appliance of modern tools of the computational fluid dynamics for the investigation of the pulsation processes in the combustion chamber caused by the design features of flame tubes and aerodynamic interaction compressor, combustor and turbine is discussed. The aim of the research is to investigate and forecast the non-stationary processes in the gas turbine combustion chambers. The results of the numerical experiments which were carried out using three-dimensional mathematical models in gaseous fuels combustion chambers reflect sufficiently the physical and chemical processes of the unsteady combustion and can be recommended to optimize the geometrical and operational parameters of the low-emission combustion chamber. The appliance of such mathematical models are reasonable for the development of new samples of combustors which operate at the lean air-fuel mixture as well as for the modernization of the existing chambers with the aim to develop the constructive measures aimed at reducing the probability of the occurrence of the pulsation combustion modes. Keywords: gas turbine engine, combustor, turbulent combustion, pulsation combustion, numerical methods, mathematical simulation.

Combustion chamber design and reaction modeling for aero turbo-shaft engine

2018 ◽  
Vol 91 (1) ◽  
pp. 94-111
Author(s):  
Raja Marudhappan ◽  
Chandrasekhar Udayagiri ◽  
Koni Hemachandra Reddy
Keyword(s):  
Numerical Modeling ◽  
Combustion Chamber ◽  
Gas Turbine ◽  
Diffusion Flame ◽  
Internal Flow ◽  
Content Type ◽  
Combustion Modeling ◽  
Flame Combustion ◽  
Aero Engine ◽  
Exit Temperature

Purpose The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance. Design/methodology/approach The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance. Findings The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments. Research limitations/implications Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities. Practical implications This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development. Originality/value The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.

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.

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