Aircraft Gas Turbine Combustor Research Articles

Development of an Augmented Conceptual Design Tool for Aircraft Gas Turbine Combustors

Combustor design is the most unreliable and challenging portion in the design process of a gas turbine. To ensure the proper performance, many experimental tests must be performed on a combustor in the industry. The above mentioned design phase is costly and time consuming. This paper focused on an automated and augmented conceptual design methodology for conventional combustors. The design tool developed for this study employs empirical and semi-empirical models which include two main parts of the combustor, the reference diameter and area as well as the component design. The necessity of this work arose from an urgent need for comprehensive and fast generation data in the conceptual design phase of a combustion chamber. This automated and comprehensive tool, equipped with the capacity to provide many details, has a considerable impact on the reduction of further experimental effort. Also, the said tool is equipped with a geometrical model generation section that has application in the future design phases, e.g., detail design.

항공용 가스터빈 연소기 기본 설계 프로그램 개발 : Part 1 - 연소기 크기 결정

Daesik Kim, Gyong Won Ryu, Ki Young Hwang and Seong Ki MinABSTRACTThis paper shows a general development process for aircraft gas turbine combustors. As a first step for develo-ping the preliminary combustor design program, several combustor sizing methodologies using reference area concepts are reviewed. There are three ways to determine the reference area; 1) combustion efficiency approach, 2) pressure loss approach, 3) velocity assumption approach. The current study shows the comparisons of the calculated results of combustor reference values from the pressure loss and velocity assumption approaches. Further works are required to add iterative steps in the program using more reasonable values of pressure loss and velocities, and to evaluate the sizing results using data for actual combustor performance and sizes.

항공용 가스터빈 연소기 기본 설계 프로그램 개발 : Part 2 - 공기 유량 배분

This study introduces the design methods for air flow distribution at the level of preliminary design, and reviews the typical combustion process and main functions of sub-components of aircraft gas turbine combustors. There are lots of design approaches and empirical equations introduced for air flow distributions at the combustors. It is shown that a decision on which design approaches work for the combustor development is totally dependent upon the objective of engine design, target performance, and so on. The current results suggested for preliminary air flow distributions need to be validated by combustor geometry checkups and performance evaluations for future works.

Large eddy simulation of soot evolution in an aircraft combustor

An integrated kinetics-based Large Eddy Simulation (LES) approach for soot evolution in turbulent reacting flows is applied to the simulation of a Pratt & Whitney aircraft gas turbine combustor, and the results are analyzed to provide insights into the complex interactions of the hydrodynamics, mixing, chemistry, and soot. The integrated approach includes detailed models for soot, combustion, and the unresolved interactions between soot, chemistry, and turbulence. The soot model is based on the Hybrid Method of Moments and detailed descriptions of soot aggregates and the various physical and chemical processes governing their evolution. The detailed kinetics of jet fuel oxidation and soot precursor formation is described with the Radiation Flamelet/Progress Variable model, which has been modified to account for the removal of soot precursors from the gas-phase. The unclosed filtered quantities in the soot and combustion models, such as source terms, are closed with a novel presumed subfilter PDF approach that accounts for the high subfilter spatial intermittency of soot. For the combustor simulation, the integrated approach is combined with a Lagrangian parcel method for the liquid spray and state-of-the-art unstructured LES technology for complex geometries. Two overall fuel-to-air ratios are simulated to evaluate the ability of the model to make not only absolute predictions but also quantitative predictions of trends. The Pratt & Whitney combustor is a Rich-Quench-Lean combustor in which combustion first occurs in a fuel-rich primary zone characterized by a large recirculation zone. Dilution air is then added downstream of the recirculation zone, and combustion continues in a fuel-lean secondary zone. The simulations show that large quantities of soot are formed in the fuel-rich recirculation zone, and, furthermore, the overall fuel-to-air ratio dictates both the dominant soot growth process and the location of maximum soot volume fraction. At the higher fuel-to-air ratio, the maximum soot volume fraction is found inside the recirculation zone; at the lower fuel-to-air ratio, turbulent fluctuations in the mixture fraction promote the oxidation of soot inside the recirculation zone and suppress the accumulation of a large soot volume fraction. Downstream, soot exits the combustor in intermittent fuel-rich pockets that are not mixed during the injection of dilution air and subsequent secondary fuel-lean combustion. At the higher fuel-to-air ratio, the frequency of these fuel-rich pockets is increased, leading to higher soot emissions from the combustor. Quantitatively, the soot emissions from the combustor are overpredicted by about 50%, which is a substantial improvement over previous works utilizing RANS to predict such emissions. In addition, the ratio between the two fuel-to-air ratios predicted by LES compares very favorably with the experimental measurements. Furthermore, soot growth is dominated by an acetylene-based pathway rather than an aromatic-based pathway, which is usually the dominant mechanism in nonpremixed flames. This finding is the result of the interactions between the hydrodynamics, mixing, chemistry, and soot in the recirculation zone and the resulting residence times of soot at various mixture fractions (compositions), which are not the same in this complex recirculating flow as in nonpremixed jet flames.

The application of porous media to simulate the upstream effects of gas turbine injector swirl vanes

Numerical methods are invaluable in the initial phase of a project for evaluating potential design solutions. The design of aircraft gas turbine combustor external aerodynamics is one example where such methodology is applied. The system is geometrically complex and requires a high fidelity mesh to resolve small geometric features, such as the injector swirl vanes. Typically the injector alone accounts for approximately 20% of the mesh cells of the entire domain. This makes it expensive to model in terms of both requisite computational resource and run time. This paper presents a numerical method for replacing the injector swirl vane passages with mathematically porous volumes which replicate the required pressure drop. This is preferential to modelling the fully featured injector as it allows a significant reduction in the cell count. This in turn allows significant time savings that serve to expedite the initial design of the combustor external aerodynamics. It also makes the geometry easier to change, scale and re-mesh during development. This method has been rigorously tested through a range of approach conditions to ensure that it is robust enough for use in the design process. The loss in accuracy owing to the simplification has been demonstrated to be less than 4.4%, for all tested flowfields. This error is dependent on the flow conditions and is generally much less for passages fed with representative levels of upstream distortion.

1118 航空用ガスタービンサブスケール燃焼器内の旋回流れを伴った噴霧燃焼におけるNO_x生成特性に関する研究(GS-6,21 ガスタービン)

1118 航空用ガスタービンサブスケール燃焼器内の旋回流れを伴った噴霧燃焼におけるNO_x生成特性に関する研究(GS-6,21 ガスタービン)

Turbulence effects on evaporation rate-controlled spray combustor performance

We consider here the effects of gas-phase turbulence on the evaporation dynamics of a polydispersed dilute fuel spray in an adiabatic well-stirred reactor (WSR) supplied with hot compressed air. Turbulence increases the time-averaged rate of heat diffusion-controlled droplet evaporation but the augmentation factor is droplet size-dependent because of droplet inertia and evaporative drag reduction. Consequently, each droplet in a spray population will not be characterized by the same rate of area change even in the same time-averaged environment. When these physical phenomena are convoluted with the residence time distribution characterizing a WSR (simulating, say, the primary zone of an aircraft gas turbine combustor), we predict the resulting fraction evaporated, evaporation rate-controlled combustion “intensity”, and corresponding exit droplet size distributions (DSD).

Electrically charged small soot particles in the exhaust of an aircraft gas-turbine engine combustor: comparison of model and experiment

Abstract The emission of electrically charged soot particles by an aircraft gas-turbine combustor is investigated using a theoretical model. Particular emphasis is placed on the influence of the fuel sulfur content (FSC). The model considers the production of primary “combustion” electrons and ions in the flame zone and their following interaction with molecular oxygen, sulfur-bearing molecules (e.g. O 2 , SO 2 , SO 3 , etc.) and soot particles. The soot particle size distribution is approximated by two different populations of mono-dispersed large and small soot particles with diameters of 20–30 and 5–7 nm, respectively. The effect of thermal ionization of soot and its interaction with electrons and positive and negative ions is included in the model. The computed positive and negative chemiion (CI) concentrations at the combustor exit and relative fractions of small neutral and charged soot particles were found to be in satisfactory agreement with experimental data. The results show that the FSC indeed may influence the concentration of negative CI at low fuel flow into combustor. Importantly the simulation indicates a very efficient mutual interaction of electrons and ions with soot particles with a large effect on both ion and charged soot particle concentrations. This result may be interpreted as a possible indirect effect of FSC on the growth and size distribution of soot particles.

Gaseous SO3 and H2SO4 in the exhaust of an aircraft gas turbine engine: measurements by CIMS and implications for fuel sulfur conversion to sulfur (VI) and conversion of SO3 to H2SO4

Abstract An estimation of the conversion efficiency ( e ) of fuel sulfur to SO 3 and H 2 SO 4 , where e =([SO 3 ]+[H 2 SO 4 ])/[S T ] and [S T ] is the total sulfur atom concentration in the exhaust at the exit of an aircraft gas-turbine combustor, was derived from measurements by comparison with model results. The major results of the presented CIMS experiments and their interpretation with a model simulation are: (i) The efficiency is e =2.3±1% at an exhaust age of about 5 ms from the combustor exit; (ii) The SO 3 molecules represent a major fraction of sulfur (VI) gases e A 3 -conversion to H 2 SO 4 takes place in the sampling line where the exhaust gases spend a sufficiently long time and where the temperature is lower than in the hot exhaust. The coincidence of e from our work (measurements with the sampling point in the exhaust just behind the combustor exit) and e the measurements in an exhaust at a plume age of about 1 s suggests that the sulfur (VI) formation is inefficient in the post-combustor flow inside the aircraft engine.

Design and Development of a Research Combustor for Lean Blow-Out Studies

In a modern aircraft gas turbine combustor, the phenomenon of lean blow-out (LBO) is of major concern. To understand the physical processes involved in LBO, a research combustor was designed and developed specifically to reproduce recirculation patterns and LBO processes that occur in a real gas turbine combustor. A total of eight leading design criteria were established for the research combustor. This paper discusses the combustor design constraints, aerothermochemical design, choice of combustor configurations, combustor sizing, mechanical design, combustor light-off, and combustor acoustic considerations that went into the final design and fabrication. Tests on this combustor reveal a complex sequence of events such as flame lift-off, intermittency, and onset of axial flame instability leading to lean blowout. The combustor operates satisfactorily and is yielding benchmark quality data for validating and refining computer models for predicting LBO in real engine combustors.

Effects of Design Parameters on Cooling Air Requirement in a Gas Turbine Combustor

T purpose of this paper is to show the effects of design parameters on the cooling air requirement in a typical can-annular aircraft gas turbine combustor. Works in this field have been reported earlier, for instance by Whittaker, ] showing how wall temperatures in different combustion zones depend upon various parameters. The objective of this work is the opposite, i.e., to prescribe a maximum allowable wall temperature and to calculate the total amount of cooling air required in order not to exceed this temperature at any location.

Impact of Broad Specification Fuels on Aircraft Gas Turbine Combustors

An analytical study was conducted to assess the impact of the use of broad specification fuels with reduced hydrogen content on the design, performance, durability, emissions, and operational characteristics of combustors for commercial aircraft gas turbine engines. The study was directed at defining necessary design revisions to combustors designed for use of Jet A when such are operated on ERBS (Experimental Referee Broad Specification) fuel which has a nominal hydrogen content of 12.8% in current Jet A. The results indicate that improvements in combustor liner cooling, and/or materials, and methods of fuel atomization will be required if the chemical and physical fuel properties approach those of ERBS.

The NOx Emission Levels of Unconventional Fuels for Gas Turbines

An experimentally verified NOx emission model for gas turbines has been reported previously. The model has been modified to determine the NOx emission levels of various fuels as compared to No. 2 distillate oil. The NOx emission levels of various conventional and unconventional gas turbine fuels of interest are predicted. The predicted NOx emission levels for these fuels, including methanol, ethanol, propane, and hydrogen, are in good agreement with available laboratory and field data from stationary, aircraft, and automotive gas turbine combustors. The predicted results should be applicable to other fuel-lean, heterogeneous combustion systems.

The Pulsed Thermocouple for Gas Turbine Application

A mechanically pulsed suction thermocouple has been developed. The gas to be measured is sucked through a sonic orifice, thus eliminating the influence of the velocity inside the combustor. The signal from the thermocouple is processed by an analogue circuit. Contrary to the usual approach to the problem of dynamic temperature measurements (i.e., the attempt to find an exact solution to the extrapolation of the temperature rise curve) in this work, a calibration of the probe was done. This calibration showed very little scatter and a good repeatability. The overall measurement accuracy was better than ±1 percent. As a test of application, a partial survey of the temperature distribution within an aircraft gas turbine combustor was done. A satisfactory agreement was observed between temperatures measured by the thermocouple and those determined from gas analysis. In this test the pulse thermocouple proved to be a reliable and fast tool for the measurement of local gas temperatures.

Gas Turbine Combustor Stabilization by Heat Recirculation

The feasibility of heat recirculation for stabilization of lean mixtures and emission reduction has been studied in detail for a typical aircraft gas turbine combustor. Thermodynamic calculations have indicated temperature and heat recirculation rates for operation of the combustor over a range of combustion zone equivalence ratios and for varying modes of desired engine operation. Calculations indicate the feasibility of stabilizing the combustion zone at equivalence ratios as low as 0.2 with achievable heat recirculation rates. Detailed chemical kinetic calculations suggest that combustor heat release is maintained with reaction completion substantially before the NO forming reactions, even though CO is rapidly oxidized in this same region.

航空用ガス・タービン燃焼器のライナにおける伝熱と冷却法

航空用ガス・タービン燃焼器のライナにおける伝熱と冷却法

The Combustion of Liquid-Fuel Spray

Abstract This paper is concerned with the evaporation and combustion of liquid-fuel spray in the range of sizes predominantly used in aircraft gas-turbine combustors. Spray of uniform-size fuel drops of 70 to 130 microns diameter was produced by means of a spinning disk. Photographic techniques were developed to obtain an indication of the rate of change of diameter and the velocity of the burning drops, while moving freely in air. Rates of burning for several pure hydrocarbon fuels were determined. The experimental work emphasizes the complexity of the combustion of a fuel spray and the need for much additional work.