As a prospective pressure gain combustion technology, detonation combustion has obvious potential for greatly increasing the thermodynamic performance of marine gas turbine due to its advantage in low entropy generation, fast heat release and self-pressurization. In this paper, a thermodynamic cycle model of detonation combustion based marine gas turbine is established considering the variable specific heat capacity. On this basis, a comparative analysis is investigated to discuss the effects of different factors on the performance enhancement of marine gas turbine by using detonation combustion. The results demonstrate that compared to the conventional deflagration combustion, detonation combustion can significantly improve the thermodynamic performance of marine gas turbine under various condition. As far as the present study is concerned, the thermal cycle efficiency can be increased to 42.97∼46.76%. Besides, it is found that the effects of pressure ratio on performance enhancements of marine gas turbine are higher than those of atmospheric temperature and temperature ratio. When pressure ratio is ranged from 13 to 30, both thermal cycle efficiency and specific power enhancements are about 20∼27%.

In the present paper, the synergistic combination of inter-cooling with pulsed detonation combustion is analyzed concerning its contribution to NO x and CO 2 emissions. CO 2 is directly proportional to fuel burn and can, therefore, be reduced by improving specific fuel consumption and reducing engine weight and nacelle drag. A model predicting NO x generation per unit of fuel for pulsed detonation combustors, operating with jet-A fuel, is developed and integrated within Chalmers University’s gas turbine simulation tool GESTPAN. The model is constructed using CFD data obtained for different combustor inlet pressure, temperature and equivalence ratio levels. The NO x model supports the quantification of the trade-off between CO 2 and NO x emissions in a 2050 geared turbofan architecture incorporating intercooling and pulsed detonation combustion and operating at pressures and temperatures of interest in gas turbine technology for aero-engine civil applications.

In order to decrease the fuel consumption, a new flight mode is being considered for twin-engine helicopters, in which one engine is put into sleeping mode (a mode in which the gas generator is kept at a stabilized, sub-idle speed by means of an electric motor, with no combustion), while the remaining engine operates at nominal load. The restart of the engine in sleeping mode is therefore deemed critical for safety reasons. This efficient new flight mode has raised the interest in the modeling of the restart of a turboshaft engine. In this context, the initial conditions of the simulations are better known relative to a ground start, in particular the air flow through the gas generator is constant, the fuel and oil system states are known and temperatures of the casings are equal to ambient. During the restart phase of the engine, the gas generator speed is kept at constant speed until the light-up is detected by a rise in inter-turbine temperature, then the starter torque increases, accelerating the engine towards idle speed. In this paper, the modeling of the acceleration of the gas generator from light-up to idle and above idle speeds is presented. Details on the light-up process are not addressed here. The study is based on the high-fidelity aero-thermodynamic restart model that is currently being developed for a 2000 horse power, free turbine turboshaft. In this case, the term high-fidelity refers not only to the modeling of the flow path components but it also includes all the subsystems, secondary air flows and controls with a high level of detail. The physical phenomena governing the acceleration of the turboshaft engine following a restart — mainly the transient evolution of the combustion efficiency and the power loss by heat soakage — are discussed in this paper and modeling solutions are presented. The results of the simulations are compared to engine test data, highlighting that the studied phenomena have an impact on the acceleration of the turboshaft engine and that the model is able to correctly predict acceleration trends.

The Marine Corps Systems Command is evaluating alternative propulsion systems to achieve high water speed for the future USMC Amphibious Combat Vehicle (ACV). A gas turbine engine is one of the propulsion systems evaluated. The primary risk of operating a gas turbine engine in the ACV is power loss due to the ingestion of marine contaminants such as saltwater mist in the air, saltwater spray generated from the vehicle operation, green water wash caused by the operation of the vehicle in the surf zone or in rough seas, and hard particles such as sand present in the marine environment. The Auxiliary Ships and New Acquisition Support Branch (Code 425) of the Naval Surface Warfare Center, Philadelphia Division conducted a study to assist the Marine Corps Systems Command in assessing the feasibility of using a gas turbine engine as a propulsion system on future USMC ACVs. The study was focused on developing and testing a gas turbine intake solution for the ACV that can remove saltwater from the intake airstream of the notional 3,000 horsepower ACV engine. Code 425 developed a two-part solution for the intake of the ACV. The first part of the solution is the Combustion Air Protection Shroud (CAPS) located at the entrance of the engine intake and designed to protect the ACV engine from green water wash by elevating the intake above the ACV deck. The second part of the solution is a gas turbine intake filtration system located downstream of the intake shroud and designed to remove marine contaminants that enter the intake shroud. A reduced-scale model of the CAPS was designed by Code 425 in conjunction with Gibbs & Cox and tested at the Davidson Laboratory High Speed Test Basin at the Stevens Institute of Technology to determine the optimum extension height of the CAPS to protect the engine intake. This paper covers the design and testing of the CAPS. The results showed that a 2.67 ft. tall CAPS with selectively closeable air intake louvers is sufficient to keep out saltwater from the ACV gas turbine engine.

Generation of combustion noise is here studied in a realistic turbine stage configuration, where the temperature fluctuations generated by an aeronautical burner (responsible of the direct combustion noise generation, often neglected in most of the studies) are modelled using a cylindrical fluctuating energy source term placed upstream of the inlet guide vanes. This modelling not only better accounts for the actual physical combustion noise generation phenomena, but also for the actual stratification of temperature at the combustor exit and covers a broader range of exciting frequencies as in actual engines. The temperature fluctuations generated by this realistic fluctuating energy source term are then convected through the turbine stage and the resulting noise is analyzed and compared with previous LES results with ideal monochromatic entropy forcing and analytical compact methods. New physical insight on the various noise generation mechanisms and mode conversion is also provided.

A program suite of coupled analysis tools for all components of industrial gas turbines developed by the authors was applied to off-design operation of single- and two-shaft gas turbines. The through-flow methods are based on a stream function approach and a finite element solution procedure. They include high-fidelity loss and deviation models with improved correlations. Advanced radial mixing and end-wall boundary layer models were applied to simulate 3D flow effects. For air-cooled turbine analysis, various types of cooling air injection were encompassed: film cooling, trailing edge injection and disc/end-wall coolant flow. Several improvements and extensions of the meridional solvers involved were employed. For axial compressor analysis, the original perfect gas treatment was replaced with the same real gas library that had already been implemented for flow computations in air-cooled expansion turbines and combustion modules. Ambient conditions with wet inflow of the compressor can now be taken into account. Compared to the previous model, better closure of the thermodynamic cycle can be achieved with the new approach. The meridional flow solver applied to the cooled turbine was upgraded and has become more stable for transonic outlet Mach numbers. With regard to the cooling flow supply, the codes offer greater freedom at the compressor extraction and turbine injection stations, allowing more realistic modeling of the secondary flow bifurcation. Off-design operating conditions and sensitivity studies for both single- and two-shaft gas turbines covering the entire range of interest are presented and are compared with experimental rig-engine data where available.

The development of new generations of aircraft engines with reduced environmental impact heavily relies on high-fidelity 3D numerical analysis of the main engine components, compressor, combustion chamber, turbine and their interactions, including the transient and off-design behavior of the full engine. Unlike component-by-component analysis, which requires separate assumptions for the pressure and temperature boundary conditions for each component, a fully coupled approach requires only knowledge of the compressor inlet and turbine outlet flow conditions. In addition, the engine rotation speed can also be varied during the simulation to converge to the correct balance of power between compressor and turbine. This integrated approach provides a detailed description of the flow field inside the full engine at the desired operating point with one single CFD simulation. The full engine simulation methodology can be developed at several levels: (1) RANS simulations with mixing-plane interfaces between components; (2) advanced RANS treatment with inputs from the nonlinear harmonic (NLH) methodology to allow for tangential non-uniformity, such as hot streaks entering the turbine nozzle from the combustor; (3) inclusion of the unsteady rotor-stator interactions, via NLH, in compressor and turbine stages; (4) coupling with LES simulations in the combustor. This paper presents results from levels (1) and (2) of this methodology applied to a micro-turbine gas engine including the HP compressor, combustor, HP and LP turbines and the exhaust hood. The geometry has been obtained from the redesign of the KJ66 micro gas turbine engine using preliminary design tools. The injection and burning of fuel inside the combustion chamber are modeled with a simplified flamelet model. The paper presents the approach and results of the full engine simulation; as well as the initial steps towards level (3).

The objective of this study was to investigate the effects of oscillations in the main flow on film cooling at various single frequencies at a low average blowing ratio of M = 0.5. The oscillations in the main flow could be a result of combustion instabilities that have been one of the major concerns for gas turbine industry. Understanding the effect of the instabilities on film cooling is important for better design of the gas turbine engines. The frequencies from 268 to 2144 Hz were identified as the dominant frequencies from a Fourier analysis of a combustor instability data on pressure oscillations. Lastly, the experimental data on gas turbine combustor instabilities is applied to the main flow using Fourier Series and the results are compared to those at single frequencies. Numerical simulations are carried out using LES Smagorinsky-Lilly and URANS k-epsilon models. This study is focused on film cooling effectiveness and heat transfer coefficient which are very important in calculation of the blade temperatures. The results show that as the frequency of the main flow goes from 0 to 180 Hz, the film cooling effectiveness is decreased due to enhancement of jet lift off with increasing frequency. However, when the frequency goes from 180 to 268 Hz, the film cooling effectiveness climbs up sharply because a thin coolant film near the wall is overlapped by large vortices containing the coolant. If the frequency changes from 268 to 1072 Hz, the effectiveness drops because the large vortices generated catch up with each other and they start overlapping and they are moved away from the wall. Main flow frequencies from 1072 to 2144 Hz cause an increase of the film cooling effectiveness since the coolant jet could not respond to these very high frequencies and the coolant behavior starts to return to that at 0 Hz gradually along with the effectiveness. In terms of heat transfer coefficients, when the oscillation frequency climbs from 0 to 536 Hz, the spanwise-averaged Stanton number ratio (St m /St o ) increases due to growing disturbances in the flow. If the frequency is increased from 536 to 2144, the spanwise-averaged Stanton number ratio is decreased. When the oscillation frequency exceeds 536 Hz, the mixing between the hot mainstream and the coolant is reduced because jets do not respond to the flow oscillations as quickly by very short period.

Material selection for key components is a critical step in the design process where high temperature and large loads are involved. Service conditions, however, may be more aggressive than those to which many materials have been exposed to in laboratory conditions. For example, key combustion equipment will experience super-imposed thermal, mechanical, and vibratory loading. Despite the more recent efforts to characterize materials under these so-called combined extreme environments (CEEs), the temperature- and rate-dependent cyclic hardening/softening responses are still relatively under-characterized. A method to develop first approximations of constitutive model/parameters for materials under extreme service conditions is presented. The method is exercised on IN617 a Ni-base alloy often employed for high temperature applications. The approach shows that a minimal collection of tensile, creep, and LCF data are needed to develop predictions of materials under thermomechanical fatigue with vibratory loading.

The emission laws for internal combustion engines become more and more strict. Therefore, new concepts have to be implemented. In the so called Miller approach the intake valve is closed before the intake stroke is finished, thus resulting in a lower combustion end temperature and pressure. As a negative result, the specific power of the engine is reduced. This disadvantage has to be compensated by an increased boost pressure delivered by a turbocharger compressor. For the turbomachinery this means for low end torque engine operation a compressor operating point at high pressure ratio and low mass flow. Thus an increased risk for surge results. A cost-effective measure to establish an utilizable compressor map in this regime is a ported shroud casing treatment. Here, a circumferential cavity connects the low channel at the inducer with the compressor housing inflow, allowing fluid to recirculate at low mass flows. Thereby, the gross inducer mass flow is increased, the flow stabilized and hence the surge line improved. In this paper, a ported shroud casing treatment is developed employing CFD. The aim is to improve the surge line as well as the stability of the compressor characteristics and to minimize the impact on compressor efficiencies at high flows as well as the acoustic behaviour at the same time. In order to validate the performance of the design, standard hot gas measurements as well as acoustic measurements are conducted and analyzed. Furthermore, the impact of a commonly applied 90° inflow bend on the performance of the ported shroud cavity is investigated by experimental data.

In this study, research activity concerning the utilization of biogas into a commercial micro gas turbine (MGT) Ansaldo Turbec T100P is presented. The use of biogas in MGT has been already quite explored in literature. However, conflicting and not yet fully comprehensive results highlight the need for further scientific investigations. Although several authors have demonstrated, through numerical studies, the possibility of using biogas with amounts of CO 2 up to 40–50% in volume in commercial MGT, experimental evidences of this condition are not yet present in literature. In fact, the CO 2 concentration limit experimentally validated in literature, after which phenomena of detachment and instability of the flame happen, is considerably lower (∼12%). The utilization of biogas in micro gas turbines generally requires structural adaptations, often invasive and expensive, to the fuel compression unit and control valves. Not infrequently a partial redesign of the combustor is often necessary. Costs associated with the modifications significantly reduce the benefits of a rational and sustainable exploitation of energy resources of lower value as the biogas. Aim of this activity is to overcome these issues, by identifying a non-invasive and cost-effective technical solution to increase the CO 2 concentration in the fuel and to extend the operation domain when low calorific value fuels are used. It has been studied, implemented and tested an innovative management strategy of fuel feeding and combustion. Experimental layout and preliminary results of this novel approach are shown and discussed.

The purpose of the paper is the investigation of the combustion development in a Lean Premixed Prevaporized (LPP) combustor supplied with liquid fuels of a Micro Gas Turbine. The injection system is equipped by an airblast atomizer with three air inlets. In such a study the analysis is performed by a CFD tool that can simulate the injection conditions, by isolating and studying some specific phenomena. A 3-D fluid dynamic code (the FLUENT ® flow solver) has been used to simulate the spray pattern in the chamber fuelled by kerosene fuel. Preliminarily, the numerical simulation refers to cold flow conditions, in order to validate the estimation of the fundamental spray parameters, such as the spray angle and Sauter Mean Diameter of the droplets; in a second step, the calculations employ boundary conditions close to those occurring in the actual combustor operation, in order to predict the fuel vapour distribution throughout the premixing chamber. In particular, the fuel is injected under the typical conditions that occur in the injection system of a gas turbine LPP combustor and in all cases examined, the boundary information is introduced in terms of air and fuel mass flow rates and of inlet characteristics of the air flow entering the revalorizing chamber, in order to estimate the fuel vapour formation and its distribution. In a third phase the combustion phenomenon is simulated and the NO emissions calculated. Finally, the best solution obtained is also tested by using a bio-fuel to compare the combustion performance and NO amount.

Ammonia is expected to be a hydrogen carrier that has potential as a carbon-free fuel. Ammonia is known as a nonignitable fuel, and it is not easy to hold ammonia flames under atmospheric conditions. A demonstration test with the aim of showing the potential of ammonia-fired power plants was conducted using a micro gas turbine. A 50-kW-class turbine system firing kerosene was selected as a base model. More than 40 kW of power generation was achieved by firing ammonia gas or a mixture of ammonia and methane by modifying the combustor, the fuel control device, and the gas turbine startup sequence. The prototype bifuel combustor is a swirl combustor employing a non-premixed flame and a decreased air flow rate near a gas fuel injector for flame holding. Ammonia combustion in the prototype bifuel combustor was enhanced by supplying hot combustion air and by modifying the air inlets. However, the exhaust gases from the ammonia flames had high NOx concentrations. NOx removal equipment using selective catalytic reduction can reduce NOx emission levels to below 10 ppm from more than 1000 ppm (converted value of NOx to 15% O 2 ) as already reported. However, downsizing of NOx removal equipment should be achieved for practical use. Therefore, a low NOx combustor was developed. As the first step of the development of the combustor, flame observation in the gas turbine combustor was tried. Although the observation area was limited, an inhomogeneous swirling orange flame of ammonia gas was observed. Then, a combustor test rig was prepared for a detailed observation of ammonia flame under various combustion conditions. The combustor test rig used a regenerative heat exchanger for heating combustion air, and it used an orifice for pressure drop instead of a turbine. Combustion air and cooling air were supplied from two air compressors. At startup of the combustor test rig, a spark plug was used to ignite non-premixed methane and air. After heating the regenerative heat exchanger, ammonia gas was supplied to the combustor instead of methane gas. The exhaust gases from the combustor were analyzed using FTIR (Fourier transform infrared spectroscopy) under various conditions, such as methane firing, methane–ammonia firing, and ammonia firing. Although there are several concepts for NOx reduction, a rich–lean combustion method was applied first for ammonia firing. The rich–lean combustor modified from the prototype bifuel combustor also could burn ammonia well in cases of both methane–ammonia firing and ammonia firing. The rich–lean combustor succeeded in reducing NOx emission from methane–ammonia combustion to half the value measured in the case of the prototype bifuel combustor.

As a means of meeting ever increasing emissions and fuel economy demands car manufacturers are using aggressive engine downsizing. To maintain the power output of the engine turbocharging is typically used. Compared to Mono scroll turbines, with a multi-entry system the individual volute sizing can be better matched to the single mass flow pulse from the engine cylinders. The exhaust pulse energy can be better utilised by the turbocharger turbine improving turbocharger response. Additionally the interaction of the engine exhaust pulses can be better avoided, improving the scavenging of the engine. Besides the thermodynamic advantages, the multi-entry turbine represents a challenge to the structural dynamic design of the turbine. A higher number of turbine wheel resonance points can be expected during operation. In addition, the increased use of exhaust pulse energy leads to a distinct accentuation of the blade vibration excitation. Using validated engine models, the interaction of the multi-entry turbine with the engine has been analyzed and various operating points, which may be critical for the blade vibration excitation, have been classified. These operating points deliver the input variables for unsteady computational flow dynamics (CFD) analyses. From these calculations unsteady blade forces were derived providing the necessary boundary conditions for the structural dynamic analyses by spatially and temporally high-resolved absolute pressures on the turbine surface. Goal of the investigation is to identify critical operating conditions. Important is also to investigate the effect of a scroll connection valve on blade excitation. The investigations utilize validated tools that were introduced and successfully applied to several turbine types in a series of publications over recent years. It can be stated that the engine operating condition and the admission type significantly influence the forced response reaction of the blade to the different excitation orders (EO). In case of equal admission even (or multiples of two) EOs generate the largest dynamic blade stress as can be expected due to the two turbine inlet segments. This reaction also increases with the engine speed. In the case of unequal admission, the odd EOs produce the largest forced response reaction. The maximum dynamic blade stress occurs in the region where the scroll connection is just closed. Above all, the scroll connection valve influences the Beta value and thus the basic behavior — unequal or equal admission. It has been possible to reconstruct the forced response behavior of the turbine blade within an engine combustion cycle. For the first time it could be shown for a double scroll application that there is a significant dynamic blade stress change dependent on the engine crankshaft angle. Certainly, due to the inertia of the mass and damping (mass, structure, flow), the blade will not exactly follow the predicted course. However, it is clear that the transient processes within an engine combustion cycle will affect the dynamic blade stress. This applies to the turbine wheels investigated in the work at hand with low damping, high eigenfrequencies and the considered internal combustion engines — as they are typically used in the passenger car sector.

Lean burn combustion is increasing its popularity in the aeronautical framework due to its potential in reducing drastically pollutant emissions (NO x and soot in particular). Its implementation, however, involves significant issues related to the increased amount of air dedicated to the combustion process, demanding the redesign of injection and cooling systems. A reduced coolant mass flow rate in conjunction with higher compressor discharge temperature negatively affect the cooling potential thus requiring the exploitation of efficient schemes such as effusion cooling. This work describes the experimental and numerical final validation of an aeronautical effusion-cooled lean-burn combustor. Full annular tests were carried out to measure temperature profiles and metal temperature distributions at different operating conditions of the ICAO cycle. Such an outcome was obtained also with an in-house developed CHT methodology (THERM3D). RANS simulations with the Flamelet Generated Manifold combustion model were performed to estimate aerothermal field and heat loads, while the coupling with a thermal conduction solver returns the most updated wall temperature. The heat sink within the perforation is treated with a 0D correlative model that calculates the heat pickup and the temperature rise of coolant. The results highlight an overall good capability of the proposed approach to estimate the metal temperature distribution at different operating conditions. It is also shown how more advanced scale-resolving simulations could significantly improve the prediction of turbulent mixing and heat loads.

The sCO 2 power cycle concept is identified as a potentially efficient, economical, and pollutant free power generation technique for future power generation. Recent work in the literature provides some strategies and best operating conditions for direct-fired sCO 2 combustors based on zero-dimensional reactor modeling analysis, however there is a need for a detailed investigation using accurate combustion chemical kinetics and thermophysical models. Here, the sCO 2 combustor is modelled by coupling perfectly stirred reactor (PSR) and plug flow reactor (PFR) models. The real gas effects are incorporated using the Soave-Redlich-Kwong (SRK) equation of state. Also, the detailed Aramco 2.0 kinetic mechanism is used for the combustion kinetic rates. It is found that the primary zone must be diluted either with thirty or forty-five percent of the total CO 2 in the cycle to have a feasible combustor design. However, the forty-five percent dilution level at 950 K and 1000 K yielded a better consumption of CO, O 2 and CH 4 . Also, the cross-sectional area of the sCO 2 combustor can be scaled-down to 10 to 20 times smaller than a traditional combustor with the same power output. Further, from this investigation, it is also recommended to have a gradually increasing secondary dilution in the dilution zone, by using progressively larger diameter holes. This design would help retain relatively high temperature in the initial portion of the dilution zone and would help consume fuel species such as, CO and CH 4 . It appears that, for sCO 2 combustors “lean burn” is the better strategy over stoichiometric burning to eliminate CO build up at the combustor exit. The lean burn condition at equivalence ratio (ϕ) equal to 0.9 is recommended for sCO 2 combustor operation. Also, the length of the dilution zone can be scaled-down to 50% by lean burn operation of the combustor. It is also observed that the lean burn increases the net turbine power. Current work provides crucial design considerations for the development of advanced sCO 2 combustors to be used with direct-fired power cycles.

The reactor residence time required for a sCO 2 combustor is comparatively higher than an equal power, airdiluted conventional combustor. Therefore, the strategies to reduce the reactor residence time are very important in the design of sCO 2 combustors. The current work recommends a method to reduce the residence time requirement in the sCO 2 combustion chamber. Here, the combustor is modelled by coupling the perfectly-stirred-reactor (PSR) and plug-flow-reactor (PFR) models along with the detailed Aramco 2.0 combustion chemical kinetic mechanism. The real gas effects are considered by using the Soave-Redlich-Kwong (SRK) equation of state incorporated in CHMEKIN-RG. Though, the CO emission level at the exit of the primary zone of the sCO 2 combustor is lower or in some cases equal to the conventional combustor, the further decline of CO in the dilution zone is identified as very poor. Therefore, very high CO levels can be expected at the exit of the sCO 2 combustor compared to conventional combustors. CO from the sCO 2 combustor exhaust can be eliminated by lean operation of the combustor and the excess O 2 retained in the re-cycled CO 2 stream due to lean operation can be mixed with primary methane before entering the primary combustion zone. This strategy is shown to reduce the primary zone residence time requirement of sCO 2 combustion. However, the minimum level of O 2 in the re-cycled CO 2 stream is approximately 5000 ppm and the minimum required residence time in this pre-mixing chamber is around 4 ms. Also, it is observed that the primary zone residence time requirement decreases linearly with respect to the O 2 level in the re-cycled CO 2 stream.

Sandia National Laboratories (SNL) has recently purchased a supercritical carbon dioxide (sCO 2 ) turbocompressor that operates at 118,000 rpm, 750 °C turbine inlet temperature, and 42.9 MPa compressor discharge pressure, and is sized to pressurize the flow for a 1 MWe closed Brayton cycle. The turbocompressor is a line replaceable unit designed by Peregrine Turbine Technologies (PTT) located in Wiscasset, Maine, as part of their closed Brayton electric power genset rated at 1 MWe. Both this machine and a 6MW variant are intended for commercial applications burning a variety of aircombustible fuels including biomass materials. Sandia purchased this turbocompressor as the first phase of a program to construct a 1 MWe commercially viable sCO 2 recompression closed Brayton-cycle system. During this phase, the development platform resident at the SNL Brayton Lab was reconfigured to support testing of the PTT turbocompressor to moderate, or idle, conditions. The testing infrastructure at the Brayton Lab limited maximum pressure to 13.8 MPa. This pressure limitation consequently limited turbocompressor operations to a speed of 52,000 rpm and a turbine inlet temperature of 150 °C. While these conditions are far removed from the machine design point, they are sufficient to demonstrate a range of important features. Numerous testing objectives were identified and researched, most notably: the development of a reliable cycle bootstrapping process for a motorless turbocompressor; the demonstration of consistent start, steady state, and shutdown performance and operations; performance demonstration of the numerous internal seals and bearings designs that are new to this environment; demonstration of controllability via turbine back pressuring and turbine inlet temperature; and turbomachinery performance map validation. This paper presents the design and development of the testing platform, the PTT turbocompressor and progress achieved on each of the objectives.

There has been some recent work on the global kinetic modeling of flames in oxy-fuel combustion for methane. The main challenge is that none of the mechanisms were developed to understand the time-scales of ignition. Here, a 3-step mechanism was developed for methane combustion in oxy-fuel environment. The mechanisms were simulated using a closed batch homogeneous batch reactor with constant pressure and compared to baseline simulations performed using a detailed mechanism. All simulations were performed for methane used a mixture of X CH4 = 0.05, X O2 = 0.10 and X CO2 = 0.85. Mechanisms were altered using the global mechanism equilibrium approach to ensure that the steady-state values matched the reference values and were further altered using an optimization scheme to match experimental data that was taken in a shock tube. Simulation results of methane, CO time-histories, and temperature profiles from the global mechanism were compared to those from the detailed mechanism. Ignition delay times were used to represent the time-scales of combustion. This was defined for current simulations as the time required for methane concentration to reach 5% of its initial value during combustion. Using this approach, the 3-step methane combustion mechanism showed excellent improvement in the ignition timing over a range of pressures (1 to 10 bar) and initial temperatures (1500 to 2000 K) for both lean and stoichiometric mixtures but fails to predict ignition delay times at 30 bar or the ignition delay times of fuel rich mixtures. Ongoing effort focuses on extending this new global mechanism to higher pressures and to syngas mixtures.

Direct fired oxy-fuel combustion provides a promising method for heat addition into a supercritical carbon dioxide (sCO2) power cycle. Using this method of thermal energy input into the cycle allows for potentially higher fuel to bus bar cycle efficiency. In addition, the nature of the sCO2 power cycle lends itself to easy and efficient capture of 99% of the CO2 generated in the combustion process. sCO2 power cycles typically operate at pressures above 200 bar, and due to the high degree of recuperation found in these cycles, have a very high combustor inlet temperature. Past works have explored combustor inlet temperatures high enough to be in the autoignition regime. The inlet temperatures which will be explored in this work will be limited to 700°C, which will allows for very different combustor geometry than that which has been studied in the past. While this combustor inlet temperature is lower than that previously studied, when combined with the extremely high pressure, this poses several unique and difficult design challenges. In order to explore these unique design conditions a reliable and robust CFD solution method was developed. This reliable CFD solution methodology enables rapid iteration on various geometries. This paper will explore the CFD modeling setup and the assumptions which were made in the absence of well experimental data in this combustor regime. Exploration of methodology to account for possible variations in chemical kinetics due to the lack of validated kinetic models in the current literature will also be discussed. The results from the CFD runs will be discussed and the combustor design, and next steps to complete a detailed combustor design will also be discussed. This work will enable future work in the development of oxy-fuel combustors for direct fired sCO2 power. This promising technology enables the use of fossil fuels with up to 99% carbon capture, while maintaining an overall cycle efficiency competitive with natural gas combined cycle power plants.