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Authors in recent publications shared their personal “gripe” for the modeling capabilities not reasonably calibrated before the start of their application in gas turbine combustion technology programs conducted since 1973. Concurrently, the supporting fundamental research results did not provide the insight needed for making critical design decisions during the technology and/or product development phase. Therefore, a joint effort was kicked off in 2012 on fundamental research to gauge what it takes to calibrate or validate the models while in parallel support hypothesis-based technology development approach specifically targeted for developing swirl-venturi lean direct injection (SV-LDI) technology. As summarized in this paper, based on this pilot project’s output to date, for a properly formulated applicable fundamental research with limited resources, it takes much longer than what Mongia had wished for since 1973.

The empirical/analytical combustion design methodology practiced since middle 1970s comprises continuously evolving conventional design practice with significantly increased roles played by hypotheses formulation, its direct or indirect verification, semi-analytical models and multidimensional computational tools. Rapid advances in “applicable CFD” and turbulent combustion modeling during the early 1970s with remarkable contributions made by the team led by Prof. Spalding combined with resources (dollars, people, and facilities) provided by industry (Garrett, Allison, GE Aviation, Goodrich, Parker, and Woodward), government (NASA, the US Air Force, Army and Navy) and numerous universities led to formulation and successful applications of empirical/analytical design methodology in several gas turbine combustion technology and design programs. These programs included NASA staged combustion Concept 3 (in 1977), two Army combustor concepts for small engines (1978), 2100, 2400, and 2900 °F temperature rise combustors (1981–1983), the two first product combustors (1986), two near-stoichiometric temperature high-performance combustors (1993), entitlements for ultralow NOx premix/pre-vaporized and partially premixed mixers (1993, 2008), an RQL combustor for the largest turbofan engine (1996), the second-generation lean-dome combustion technology TAPS demonstration (2003) and its product introduction in GEnx (2009), NASA LDI-2 and LDI-3 technology demonstrations in 2014 and 2018, respectively. An overview of these activities along with the most recent CFD simulation and diagnostics activities are described in this chapter as a recognition of 50 Years of CFD in Engineering Sciences that has provided useful insight for advancing combustion technologies and products while simultaneously improving design process efficiency.

Combustion characteristics in a hydrogen fuelled single-element lean direct injection (LDI) combustor for low emission gas turbine engine are computationally investigated. A single-element LDI combustor was produced and simulated with steady-state Reynolds-averaged Navier–Stokes reacting computations. Reacting computations of the single-element LDI combustor were performed with an augmented reduced 23-step reaction mechanism for hydrogen/air. A realizable k − e model was used to obtain turbulence closure. For the single-element LDI combustor with different equivalence ratio, the effective area, axial velocity, total pressure drop coefficient, total temperature, mass fraction of OH, emission index of pollutant NO (EINO) were achieved and discussed.

This chapter first reviews recent research efforts directed at improving the operability range of lean direct injection (LDI) combustion technologies. Recognizing that swirl strength and flare geometry can significantly affect operability and emissions, a representative LDI swirl injector fundamental research has been undertaken to (1) investigate the impact of air swirler vane angle on the reacting flow field and flame behavior and (2) elucidate the influence of flare geometry for understanding the fundamental differences between LDI and airblast mixers. The results of these fundamental studies are discussed in order to identify design changes for improving LDI performance.

This chapter reviews and discusses the recent development in Lean Direct Injection (LDI) combustion technology. An extended definition of LDI is also provided, followed by a broad description of LDI concepts and examples. Recognizing the needs and opportunities to expand the operability range of LDI, fundamental research has been undertaken to elucidate the effects of air swirler vane angle, pressure drop, air swirler rotation direction, and overall equivalence ratio on LDI flow field and flame behavior. Additional investigation was further conducted to understand fundamental differences between representative LDI and airblast injectors. Results of these fundamental studies are discussed to help identify design changes for improving LDI performance.

An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements for next-generation LDI-3 combustor design. The National Combustion Code (NCC) was used to perform non-reacting and two-phase reacting flow computations for a nineteen-element injector array arranged in a three-module, 7-5-7 element configuration. All computations were performed with a consistent approach of mesh-optimization, spray-modeling, ignition and kinetics-modeling with the NCC. Computational predictions of the aerodynamics of the injector were used to arrive at an optimal injector design that meets effective area and fuel-air mixing criteria. LDI-3 emissions (EINOx, EICO and UHC) were compared with the previous generation LDI-2 combustor experimental data at representative engine cycle conditions.

To investigate the combustion characteristics in multi-point lean direct injection (LDI) combustors with hydrogen/air, two swirl–venturi 2 × 2 array four-point LDI combustors were designed. The four-point LDI combustor consists of injector assembly, swirl–venturi array and combustion chamber. The injector, swirler and venturi together govern the rapid mixing of hydrogen and air to form the mixture for combustion. Using clockwise swirlers and anticlockwise swirlers, the co-swirling and count-swirling swirler arrays LDI combustors were achieved. Using Reynolds-Averaged Navier–Stokes (RANS) code for steady-state reacting flow computations, the four-point LDI combustors with hydrogen/air were simulated with an 11 species and 23 lumped reaction steps H2/Air reaction mechanism. The axial velocity, turbulence kinetic energy, total pressure drop coefficient, outlet temperature, mass fraction of OH and emission of pollutant NO of four-point LDI combustors, with different equivalence ratios, are here presented and discussed. As the equivalence ratios increased, the total pressure drop coefficient became higher because of increasing heat loss. Increasing equivalence ratios also corresponded with the rise in outlet temperature of the four-point LDI combustors, as well as an increase in the emission index of NO EINO in the four-point LDI combustors. Along the axial distance, the EINO always increased and was at maximum at the exit of the dump. Along the chamber, the EINO gradually increased, maximizing at the exit of chamber. The total temperature of four-point LDI combustors with different equivalence ratios was identical to the theoretical equilibrium temperature. The EINO was an exponential function of the equivalence ratio.

Three variations of a low emissions aircraft gas turbine engine combustion concept were developed and tested. The concept is a second generation swirl-venturi lean direct injection (SV-LDI) concept. LDI is a lean-burn combustion concept in which the fuel is injected directly into the flame zone. All three variations were based on the baseline 9- point SV-LDI configuration reported previously. The three second generation SV-LDI variations are called the 5-recess configuration, the flat dome configuration, and the 9- recess configuration. These three configurations were tested in a NASA Glenn Research Center medium pressure flametube. All three second generation variations had better low power operability than the baseline 9-point configuration. All three configurations had low NO(sub x) emissions, with the 5-recess configuration generally having slightly lower NO(x) than the flat dome or 9-recess configurations. Due to the limitations of the flametube that prevented testing at pressures above 20 atm, correlation equations were developed for the at dome and 9-recess configurations so that the landing-takeoff NO(sub x) emissions could be estimated. The flat dome and 9-recess landing-takeoff NO(x) emissions are estimated to be 81-88% below the CAEP/6 standards, exceeding the project goal of 75% reduction.

The mechanism that causes a periodic combustion oscillation in a Lean Premixed Prevaporized (LPP) Combustor was experimentally investigated. The combustor was run with high flowrates of preheated air at elevated presures, with Jet-A as the fuel, and with a realistic fuel injector. The injector provides LPP combustion and features a stable diffusion flame as a pilot surrounded by a premixed main flame. The combustor was stable over a wide range of conditions, but for off-design conditions oscillations associated with unsteady flashback and liftoff of the flame base were observed. These low-frequency combustor instabilities were a strong function of a global velocity gradient (U/d), which is analogous to the U/d gradient used to characterize stability limits of Bunsen flames. PIV data quantified the location and strength of shear-layer vortex shedding. The main and pilot flames were visualized using PLIF images of formaldehyde. These images showed that for design conditions the pilot flame is very stable and that the main premixed flame exists in the mixing layer. It was seen that as the global gradient U/d decreases, the local velocity gradient in the mixing layer also decreases, making the flame more susceptible to perturbations from the vortex-shedding. It is proposed that the unsteadiness in the combustion can be explained using classic theories of premixed flame flashback and blowout that was developed for Bunsen burners.

A comprehensive study on confined swirling flows in an operational gas-turbine injector was performed by means of large-eddy simulations. The formulation was based on the Favre-filtered conservation equations and a modified Smagorinsky treatment of subgrid-scale motions. The model was then numerically solved by means of a preconditioned density-based finite-volume approach. Calculated mean velocities and turbulence properties show good agreement with experimental data obtained from the laser-Doppler velocimetry measurements. Various aspects of the swirling flow development (such as the central recirculating flow, precessing vortex core and Kelvin–Helmholtz instability) were explored in detail. Both co- and counter-rotating configurations were considered, and the effects of swirl direction on flow characteristics were examined. The flow evolution inside the injector is dictated mainly by the air delivered through the primary swirler. The impact of the secondary swirler appears to be limited.

Application of a partially calibrated National Combustion Code (NCC) for providing guidance in the design of the 3rd generation of the Lean-Direct Injection (LDI) multi-element combustion configuration (LDI-3) is summarized. NCC was used to perform non-reacting and two-phase reacting flow computations on several LDI-3 injector configurations in a single-element and a five-element injector array. All computations were performed with a consistent approach for mesh-generation, turbulence, spray simulations, ignition and chemical kinetics-modeling. Both qualitative and quantitative assessment of the computed flowfield characteristics of the several design options led to selection of an optimal injector LDI- 3 design that met all the requirements including effective area, aerodynamics and fuel-air mixing criteria. Computed LDI-3 emissions (namely, NOx, CO and UHC) will be compared with the prior generation LDI- 2 combustor experimental data at relevant engine cycle conditions.

Single-laser-shot temperature measurements at 5 kHz are performed in a laboratory scale swirl stabilized burner with an interchangeable swirler using femtosecond (fs) coherent anti-Stokes Raman scattering (CARS). Measurements were performed at various heights and radial locations within the flame in order to resolve the spatial flame structure. Temperature measurements are reported in the swirl flame between 400 and 2000 K with less than 3.0% error. In addition, the first simultaneous N2/CH4 CPP fs-CARS measurements are reported for measuring relative methane concentration in the nonreacting flow field. These results demonstrate the usefulness of CPP fs-CARS for fundamental studies of swirl stabilized flames.

An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements for next-generation LDI-2 design. The National Combustion Code (NCC) developed at NASA Glenn Research Center was used to perform reacting flow computations on an LDI-2 combustor configuration with thirteen injector elements arranged in four fuel stages. Reacting computations were performed with a consistent approach for mesh-optimization, liquid spray modeling and kinetics modeling. Computational predictions of Emissions Index (EINOx) and combustor exit temperature were compared with two sets of experimental data at medium and high-power operating conditions, for two different pressure-drop conditions in the combustor. The NCC simulations predicted the combustor exit temperature to within 1–2% of experimental data. The accuracy of the EINOx predictions from the NCC simulations was within 10% to 30% of experimental data.Copyright © 2014 by ASME

Combustion dynamics impact the design of both conventional diffusion e ame gas turbine combustors and lean premixed combustion systems. The occurrence of dynamics in diffusion e ame combustors can generally be eliminated through changes to the fueling system or operating characteristics driven primarily through empirical design know-how. Lean premixed combustors, such as industrial aeroderivative dry low-emissions combustors, have more persistent dynamics problems that are only partially ameliorated with application of empirical tools and controls. With continuing emphasis on reducing emissions and cost and increasing performance of turbopropulsion engine combustors, the design direction of e ight engines is approaching the lean premixed limit; thus, there is a need for improved tools and control strategies for combustion dynamics in these applications as well. The components of a combined analytical, experimental, and computational approach to understanding and controlling combustion dynamics are being developed. With successful implementation of such an analytical and control mechanism, further improvements in the emissions, stability, and durability characteristics of both e ight and industrial combustors will become possible.

This paper deals with the multicomponent nature of gas turbine fuels under high-pressure conditions. The study is motivated by the consideration that the droplet submodels that are currently employed in spray codes for predicting gas turbine combustor flows do not adequately incorporate the multicomponent fuel and high-pressure effects. The quasi-steady multicomponent droplet model has been employed to investigate conditions under which the vaporization behavior of a multicomponent fuel droplet can be represented by a surrogate pure fuel droplet. The physical system considered is that of a multicomponent fuel droplet undergoing quasi-steady vaporization in an environment characterized by its temperature, pressure, and composition. Using different vaporization models, such as infinite-diffusion and diffusion-limit models, the predicted vaporization history and other relevant properties of a bicomponent droplet are compared with those of a surrogate single-component fuel droplet over a range of parameters relevant to gas turbine combustors. Results indicate that for moderate and high-power operation, a suitably selected single-component (50 percent boiling point) fuel can be used to represent the vaporization behavior of a bicomponent fuel, provided one employs the diffusion-limit or effective-diffusivity model. Simulation of the bicomponent fuel by a surrogate fuel becomes increasingly better at higher pressures. In fact, the droplet vaporization behavior at higher pressures is observed to be more sensitive to droplet heating models rather than to liquid fuel composition. This can be attributed to increase in the droplet heatup time and reduction in the volatility differential between the constituent fuels at higher pressures. For ignition, lean blowout and idle operations, characterized by low pressure and temperature ambient, the multicomponent fuel evaporation cannot be simulated by a single-component fuel. The validity of a quasi-steady high-pressure droplet vaporization model has also been examined. The model includes the nonideal gas behavior, liquid-phase solubility of gases, and variable thermo-transport properties including their dependence on pressure. Predictions of the high-pressure droplet model show good agreement with the available experimental data over a wide range of pressures, implying that quasi-steady vaporization model can be used at pressures up to the fuel critical pressure.

A low-NO (sub x) aircraft gas turbine engine combustion concept was developed and tested. The concept is a second generation swirl-venturi lean direct injection (SV-LDI) concept. LDI is a lean-burn combustion concept in which the fuel is injected directly into the flame zone. Three second generation SV-LDI configurations were developed. All three were based on the baseline 9-point SV-LDI configuration reported previously. These second generation configurations had better low power operability than the baseline 9-point configuration. Two of these second generation configurations were tested in a NASA Glenn Research Center flametube; these two configurations are called the at dome and 5-recess configurations. Results show that the 5-recess configuration generally had lower NO (sub x) emissions than the flat dome configuration. Correlation equations were developed for the flat dome configuration so that the landing-takeoff NO (sub x) emissions could be estimated. The flat dome landing-takeoff NO (sub x) is estimated to be 87-88 percent below the CAEP/6 standards, exceeding the ERA project goal of 75 percent reduction.

This article gives an overview of the current rich-dome combustion system design, requirements and challenges; followed by the first alternative to rich-domes that have been successfully introduced as products; the lowest levels of achievable NOx (so called entitlement) as determined from small scale rig testing; summary of recent engine emissions data with “green” alternative fuels; description of the 2nd alternative to rich-dome products that may be of interest to the OEM’s for the N+2 and N+3 generation aviation engines; a brief discussion on the modeling and correlation accuracy expectations from future efforts in this area; the 3rd alternative to rich domes which was shown promising for autoignition times closer to 0.2 ms. The article concludes with a short section on operability and dynamics. Several large low-NOx rich domes’ takeoff NOx emission index is reproduced well by a simple correlation NO x RD −L = 0.0303PR 1.9722 w / R 2 = 0.9906 including Talon II and Trent1000. However, the LTO NOx is correlated well by a similarly good quality curve only for the group of combustors without Trent1000, DP / F 00 = 0.6793PR 1.2241 compared to lower value expression for Trent1000 alone given by DP / F 00 = 0.1292PR 1.6327. Consistent with the NOx stringency pattern set by CAEP4, CAEP6 and CAEP8 and longtime goal for achieving 85 % reduction in takeoff NOx at 30OPR, we propose the long-term LTO regulatory standard of CAEP / 18 = −37.763 + 2π effective December 31, 2033. The combustor inlet temperatures for desired overall pressure ratio at sea-level standard day static condition can be estimated by using T 3,SSS = 317.544OPR 0.2704; 320.1955OPR 0.272, respectively for the N+1 and N+2 generation engines. This along with generally accepted requirements for combustor operability, we have to manage significantly increased range of P3, T3 and fuel/air ratio, viz. T3: 216.67–1084 K; P3: 0.33–60 atm; ΔP: ~ (0.1–1.2) ΔPdesign; and FAR: FARmin - FARmax. The numerical values of the dome design pressure drop (ΔPdesign), minimum and maximum fuel air ratios (FARmin and FARmax) depend upon the combustion system design and its potential applications. We will assume typical values of these variables, respectively, 3–5 %, 0.005–008 and 0.025–0.040 for the aviation engines. The 1st generation of lean dome products met their original objectives of achieving lower NOx within the specified design constraints including cooling technology. They were immediately followed by the 2nd generation lean dome products known popularly as TAPS in GEnx in addition to planned LEAP-X and GE9X. The takeoff NOx of GEnx is given by NOx GEnx = 1.079 × 10-5 OPR 3.971 W / R 2 = 0.991. These products will be able to meet the proposed long-term LTO NOx regulatory standard within the generally accepted design modification and refinement process. TALON-X and recently introduced PW an interesting technologies competition. The 2nd generation lean domes produce an order of magnitude lower exhaust smoke number than the rich domes. For all other design requirements both the lean and rich domes have comparable characteristics. The effect of FT fuel blends on combustion efficiency and NOx is insignificant; but its benefits in regard to particulate emissions are enormous in terms of both the number density and mass emissions due primarily to significantly lower aromatic and sulfur contents. Future CFD and semi-empirical model should be developed with the proposed long term accuracy goal expressed in term of the standard deviation σ goal: 3 % of takeoff NOx, 7.5 % and 15 % respectively of idle CO and HC emission indices.

It is well known that fuel preparation, its method of injection into a combustor, and its atomization characteristics have a significant impact on emissions. A simple dilute spray model, which assumes that droplet heating and vaporization occur in sequence, has been implemented in the past within computational fluid dynamics (CFD) codes at General Electric (GE) and has been used extensively for combustion applications. This spray model coupled with an appropriate combustion model makes reasonable predictions of the combustor pattern factor and emissions. To improve upon this predictive ability, a more advanced quasi-steady droplet vaporization model has been considered. This article describes the evaluation of this advanced model. In this new approach, droplet heating and vaporization take place simultaneously (which is more realistic). In addition, the transport properties of both the liquid and vapor phases are allowed to vary as a function of pressure, gas phase temperature, and droplet temperature. Th...

Large-eddy simulation (LES) of turbulent premixed reacting flows in a gas turbine combustor (General Electric's lean premixed dry low-NOx LM6000) has been carried out to evaluate the potential of LES for design studies of realistic hardware. A flamelet model for the premixed flame is combined with a dynamic model for the subgrid kinetic energy to simulate the propagation of the turbulent flame in this high swirl and high Reynolds number flow. Comparison of the computed results with experimental data indicate good agreement in spite of relatively coarse grid resolution employed in the LES. These results provide significant confidence that LES capability for design studies of practical interest is feasible in the near future.

Correlations based on design database, combined with multi-dimensional computational combustion dynamics (CCD) models are used in the combustion design process. However, because of limitations in the current turbulent combustion models, numerics, and boundary conditions, CCD has provided mainly qualitative trends for aerothermal performance, emissions and liner wall temperature levels and gradients. To overcome these deficiencies, hybrid modeling approaches have been proposed to analyze existing combustors. A typical hybrid modeling approach combines empirical and semianalytical correlations with CCD to give quantitatively accurate predictions of NOx, CO, HC, smoke, lean blowout, ignition, pattern factor and liner wall temperatures. An alternate approach, anchored CCD, is described in this paper. First, the models were anchored with one of the five modern turbopropulsion engine combustors. The anchored models were then run for the other four combustors. The predicted results correlated well with measured NOx, CO, HC, LBO and exit temperature quality data, demonstrating a broader applicability of the anchored method. The models were also used for designing a new combustion concept. The pre-test prediction agreed well with sector rig data from development hardware, showing the feasibility of using the anchored methodology as a design tool.

A coupled Lagrangian Monte Carlo (MC) probability density function (PDF), Eulerian computational e uid dynamics (CFD) technique is presented for calculating steady three-dimensional turbulent reacting e ow in a gas turbine combustor. PDF transport methods model turbulence ‐ combustion interactions more accurately than conventional turbulence models with an assumed-shape PDF. The PDF was over composition only. The PDF transport equation was solved using a Lagrangian particle-tracking MC method. This MC module has been coupled with CONCERT, which is a fully elliptic three-dimensional bodye tted CFD code based on pressure correction techniques. CONCERT calculates the mean velocity and mixing frequency e eld that are required by the composition PDF in the MC module, whereas the MC module computes the PDF from which the mean density e eld is extracted and supplied to CONCERT. This modeling approach was initially validated against Raman data taken in a recirculating bluff body stabilized e ame. The computed mixture fraction and its variance (as obtained from the calculated PDF ) compared very well against the corresponding measurements made along several radial lines at different axial downstream positions and along the axis. A typical single annular aircraft engine combustor was also analyzed. In this preliminary study, the e owe eld, fuel, and temperature distribution were obtained based on the assumption of fast chemistry. The solutions obtained using the present approach were compared with those obtained using a presumed-shape PDF method. The comparison of the calculated exhaust gas temperatures using these two approaches with measurements made by a thermocouple rake appeared to indicate better agreement with the PDF transport technique.

Previous studies showed that the rich-dome takeoff NOxEI technology status of small, medium and large engines is presented well by the following expression: NNNNN = 0.04 × NOO1.8957 However, the same set of engines when plotted in terms of LTO NOx versus rated takeoff thrust give a large scatter band leading one to propose several different correlations. For example, the following expression for the Trent1000 LTO NOx passes well through the CF6-80C2LEC data indicating that perhaps several rich-dome combustors have very similar LTO NOx technology capability. LLN NNNTTTTT1000 = 12.753e0.0048×F LTO NOx of the nine large and very large modern richdome engines (viz. #10 through #17) were correlated well by the following generalized expression leading to their classification into four groups defined as α1 = 15,α2 = 0,α3 = −10 and α4 = −25, namely the 1, 2, 3 and 4 groups; the latter being the lowest LTO NOx RQL engine. LLN NNNRRR = 0.175 × F + αi The corresponding expression for the 2 generation lean dome (viz. #17) is represented well by LLN NNNRTLT = 0.226 × F − 40 which is not much different from the lowest NOx RQL engine specified as LLN NNNRRR = 0.175 × F − 25 if one knew how to scale down from nominal 500kN engine to 300kN engine size, viz. #12. Therefore, the next generation 500kN size engine needs to target for LLN NNNRTLT = 0.226 × F − 80. Therefore, the N+3 and N+4 combustion system goal is to further reduce LTO NOx from the 2 generation lean dome (viz. #18 which emits ~30 g of NO2/kN at 300kN) by 2/3 to 10 g of NO2/kN. Simultaneously, it should have LTO HC and LTO CO competitive with the N+1 generation RQL engines, namely: LLN HHN+3 ≤ 37.62e−0.0115×F LLN HNN+3 ≤ 170.7e−0.006×F NIIe HHNNN+3 ≤ 24.89e0.011×F NOMENCLATURE ACT-E: Advanced combustion technology for reducing emissions ASK: Available seat km BPR: Fan bypass ratio CAA: Clean Air Act Amendments 1970 edb-12: the ICAO aircraft engine emissions databank updated January 2012 EI: Emission index, g of pollutant/kg fuel EPA: Environmental Protection Agency F: Sea-level rated takeoff thrust, KN FB: LTO specific fuel burn, kg/kN GEA: General Electric Aviation ICAO: International Civil Aviation Organization LEC: Low Emissions Combustor belonging to the richdome vintage LD: Lean dome combustors LTO: Landing takeoff M: Mach number MPG: Miles per gallon OPR: SLSS takeoff pressure ratio (overall pressure ratio) P&W: Pratt and Whitney RF: Radiative forcing RPK: Revenue passenger km RQL: Rich-Quench-Lean approach used for lowering NOx from rich-dome combustors RR: Rolls Royce SLSS: Sea level standard day static (M=0) VOC: Volatile Organic Compounds INTRODUCTION Continuing emphasis on efficiency, and concerns on global warming, local air quality and human health have brought in many changes in the gas turbines and resulting advances in combustion technolgy during the last forty years. Several innovative approaches are currently underway to further improve propulsion system efficiency which motivated the team for an academic study and initiation for the N+3 and N+4 combustion technology and attendant research activities as briefly described in the following paragraphs. Extension and calibration of ACT-E process to propulsion engines combustion products ranging in the rated takeoff thrust from 30kN to more than 500kN has been summarized in regard to the best classes of combustors for NOx, CO and HC emissions. The mission fuel burn is covered in Mongia’s Part 4 paper followed by this paper to summarize NOx, CO, HC and smoke tradeoffs. CO2, CH4, N2O, and several other fluorinecontaining halogenated substances are the most important constituents of the greenhouse gases (GHG) directly emitted by humans. It should be mentioned that the Global Warming Potentials (100-Year Time Horizon), GWP of CH4 and N2O are 21 and 310, respectively compared to CO2. As shown in Figure 1, the U.S. GHG emissions have increased by 8.7 percent from 1990 to 2011 with an average annual increase of 0.4%

This paper summarizes academically simulated takeoff and cruise operating conditions for the N+3 and N+4 generation combustion systems and comparison likewise with the N and N+1 generation aeroengines in addition to GE9X and RB3025, potential engines introduced as marketing campaign start for the next upgrade of the B777. The technological challenges are summarized in order to meet very demanding NOx, CO and HC emissions goals without adversely impacting other design requirements that have generally been taken for granted for the rich-dome aeroengines. Technology assessments follow for several approaches including axially staged combustors with lifted LDI JIC compared to late lean concepts currently being followed by the power generation OEM’s. Premix and LDI approaches involving the use of variable geometry actuations are evaluated in addition to past and ongoing activities for developing several promising LDI concepts along with status on modeling capabilities that could be used for extending to axially staged micro LDI jets to be discussed in future publications.

The certified ICAO engine emissions database can be used for conducting comparative assessment on the fuel burn advances made by the major OEM’s during the last forty years. Comprehensive comparison of the modern engines relative to the 1 generation propulsion engines of the 1970’s era, we can claim 10-25% reduction in FB for the large rich dome engines. For the medium size rich-dome engines, the corresponding reductions are 15-50%. The lowest LTO specific fuel burn FB empirical expression (viz. GEnx-1B) for large engines is given by: FFGGGG = 3.0245e−4.1037×10 −4F The corresponding LTO FB expressions for the medium size engines (FFMMMMMM ) and a broader thrust range of 60150kN ( FFS−M) of engines are: FFMMMMMM = 3.854e−0.00101F FFS−M = 4.3216e−0.00141F The following generalized fuel burn expression (FFG) is proposed for all large engines

Comprehensive assessment of the medium size rich-dome engines was conducted leading to the following emissions correlations: (1) LTO NOx = 1.129 × OPR 1.0899 with R 2 = 0.9248 Takeoff NOxEI given by (2) NOxEI = 0.0729 × OPR 1.7197 with R 2 = 0.9603 COEI idle = 396.42 NOxEI Takeoff 0.814 These correlations may be compared with the following for the CFM56 Tech Insertion: Takeoff NOxEI CFM_TI = 0.0744 × OPR 1.7151 Idle COEI CFM_TI = 396.42 Takeoff NOxEI 0.814 Idle HCEI CFM_TI = 0.1609 × Idle COEI - 3.1959 TALON II takeoff NOxEI data are reproduced well by: NOxEI TALON II = 0.0167 × OPR 2.1403 TALON II gives 10% lower NOx at 26 OPR and its NOx is comparable with the CFM_TI at 34 OPR. The CFM DAC technology is competitive with LEC’s for the low rated thrust engines. However, interaction between the two domes leads to early quenching with resultant higher idle COEI plateau. On the other hand, the 40 OPR lean DAC gave 25% higher NOx than LEC. Moreover, lean DAC (Gen-1) impacted fuel burn adversely making its likelihood to continue as product discouraging. The second generation lean dome technology initially kicked off under NASA sponsorship with significantly larger funding support from the CFMI and GE Aviation (GEA) led to successful introduction of TAPS into products (GEnx-1B and Gen-2B) with potential applications in other future GEA engines.

A comprehensive assessment of emissions characteristics of the 1st, N and N+1 generation rich-dome combustion products has been done to identify the lowest emissions products. Focus of this paper is on the large rich-dome engines with its potential application for the (N+3) and (N+4) mixers with inspirational target takeoff NOxEI of 5 at 55 OPR. A total of ten engine models of the 1st generation were selected in addition to eight recently certified large engines. After evaluating several choices for conducting comparative assessment, the following three expressions were proposed for average takeoff NOxEI, idle COEI and HCEI entitlements, respectively: NOxEI L = 0.0288 × OPR 1.991 Idle COEI L = 815.36 Takeoff NOxEI L 1.159 Idle HCEI L = 0.15 × Idle COEI L - 2.0 In regard to application of the rich-dome technology to the (N+2) cycle based (N+3) mixers, the author tentatively gives it low probability of success barring success story stemming from Lee et al. [2012].

An effort was undertaken to identify best practices for CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with swirl-venturi swirlers for next-generation LDI design. The National Combustion Code (NCC) was used to perform non-reacting and reacting computations on several LDI injector configurations in a 3×3 element array with different vane angles. Non-reacting and reacting computations were performed with single and multi-element LDI configurations, with a consistent approach to mesh-optimization, spray-modeling and kinetics-modeling with the NCC. Computational predictions of NOx emissions index were compared with experimental data. CFD values of CO were used to assess lean blow-out predictive capability for two different multi-element, LDI configurations and compared to experimental data.Copyright © 2013 by ASME

A major challenge which emerged in developing innovative approaches for significantly reducing the aviation Greenhouse Gas (GHG) emissions was the design of a compact heat exchanger. As a part of the FAA program under US Department of Transportation Contract Number DTRT57-11-P-80083, an LP Turbine Exhaust Driven (TED-T) secondary cycle was proposed for contrail mitigation as well as an increment of 4-5% pts in overall powerplant efficiency. It was identified that a compact heat exchanger is a critical component of TED-T. In this series of papers, Paper #1 presented some of the existing advanced compact heat exchanger (HEX) technologies along with parametric analysis of a baseline configuration. A 1-D steady state code was developed for heat transfer calculations. In this paper, a similar parametric study is presented for a Printed Circuit Heat Exchanger (PCHE) operating at the proposed TED-T cycle conditions.

A major challenge which emerged while looking at approaches for reduction of aviation Greenhouse Gas (GHG) emissions was the design of a core exhaust heat recovery system. As a part of the FAA program under US Department of Transportation Contract Number DTRT57-11-P-80083, an LP Turbine Exhaust Driven (TED-T) secondary cycle was proposed for contrail mitigation as well as an increment of 3% pts. in overall powerplant efficiency. The compact heat exchanger emerged as a critical component for the realization of TED-T. In this work some of the existing advanced compact heat exchanger (HEX) technologies are identified & presented. A 1-D steady state code was developed for heat transfer calculations. A parametric analysis was carried out for the baseline configuration. In this paper, a similar parametric study is presented for an off-set finned HEX operating at the proposed TED-T cycle conditions.

A major challenge which emerged while looking at approaches for reduction of aviation Greenhouse Gas (GHG) emissions was the design of a core exhaust heat recovery system. As a part of the FAA program under US Department of Transportation Contract Number DTRT57-11-P-80083, an LP Turbine Exhaust Driven (TED-T) secondary cycle was proposed for contrail mitigation as well as an increment of 3% pts. in overall powerplant efficiency [1]. The compact heat exchanger emerged as a critical component for the realization of TED-T. In this series of papers, Paper #1[2] identified & presented some of the existing advanced compact heat exchanger (HEX) technologies. A 1-D steady state code was developed for heat transfer calculations. A parametric analysis was carried out for the baseline configuration. In this paper, a similar parametric study is presented for an off-set finned HEX operating at the proposed TED-T cycle conditions.

To enhanced gas turbine combustor performance and emissions characteristics, better design methods need to be developed. In the present investigation, an emission model that simulates a detailed chemical kinetic scheme has been developed to provide the rate of reactions of the parent fuel, an intermediate hydrocarbon compound, CO, and H{sub 2}. The intermediate fuel has variable carbon and hydrogen contents depending on operating conditions, that were selected in the development effort to simulate actual operating conditions, that were selected in the development effort to simulate actual operation of rich/lean, diffusion flame, and lean combustor concepts. The developed reaction rate expressions address also the limited reaction rates that may occur in the near-wall regions of the combustor due to the admittance of radial air jets and cooling air in these regions. The validation effort included the application of the developed model to a combustor simulated by a multiple-reactor arrangement. The results indicate the accurate duplication of the calculations obtained from the detailed kinetic scheme using the developed model. This illustrates the great potential of using such a unified approach to guide the design of various types of combustor to meet the more stringent emissions and performance requirements of next-generation gas turbinemore » engines.« less

In an effort to reduce the emissions produced by gas turbine combustors to meet the expected trends of more stringent regulations, a model to calculate NOx produced by lean combustion concepts is proposed. The model simulates the combustor by a number of reactors representing the main combustion as well as pilot zones. A detailed reaction rate scheme was used to provide a fundamental basis for the derivation of reaction temperature and NOx formation expressions. The model validation involved the utilization of the data obtained for a combustor that comprised a prechamber in which fuel was vaporized and mixed with air, and a variable geometry feature to control air split between various zones. In addition to the satisfactory agreement with the measurements obtained under a wide range of operations, the model provided insight into the roles of the primary zone and pilot in forming NOx. The present investigation also shows that multidimensional calculations can provide important assistance for emissions model development. Methods of NOx reduction may be implemented in light of such information.

Innovative approaches for significantly reducing the aviation Greenhouse Gas (GHG) emission need to be evaluated based on simple calculation methodology, viz. Numerical Propulsion System Simulation (NPSS) that can be easily and reliably executed by an academic research group with guidance provided by a panel of experts from industry, FAA, NASA, DOE and DOD. Part 1 of the series of papers gives a list of the proposed advanced Brayton cycles for undertaking an extensive study under the FAA program. Part 2 of the series of papers shows that good agreement was achieved between the NPSS calculations and engine data of several engines, including the aero-engines CFM56, CF6-80C, GE90, and industrial engines LM6000 and LMS100 with and without recuperation. Several engine configurations were analyzed in Part 3 leading to selection of the most promising configurations for the follow-on conceptual design effort by potential OEM. Part 4 extended the cycles to introduce the advantage of the LP Turbine Exhaust Driven Thermal (TED-T) secondary cycle for contrail mitigation without performance penalty. In the present paper, the engine layout is simplified by using a fuel reactivity controlled internal combustion concept which when combined optimally with the turbine engine will further reduce CO2 emission.

Innovative approaches for significantly reducing the aviation Greenhouse Gas (GHG) emission need to be evaluated based on simple calculation methodology, viz. Numerical Propulsion System Simulation (NPSS) that can be easily and reliably executed by an academic research group with guidance provided by a panel of experts from industry, FAA, NASA, DOE and DOD. Part 1 of the series of papers gives a list of the proposed advanced Brayton cycles for undertaking an extensive study under the FAA program. Part 2 of the series of papers shows that good agreement was achieved between the NPSS calculations and engine data of several engines, including the aero-engines CFM56, CF6-80C, GE90, and industrial engines LM6000 and LMS100 with and without recuperation. Several engine configurations were analyzed in Part 3 leading to selection of the most promising configurations for the follow-on conceptual design effort by potential OEM. In the present paper, extension of the advanced Brayton cycle is considered by introducing an LP Turbine Exhaust Driven Thermal (TED-T) cycle to provide heat source for a closed supercritical CO2 cycle for running either a propulsion system or an electric generator.

Innovative approaches for significantly reducing the aviation Greenhouse Gas (GHG) emission need to be evaluated based on simple calculation methodology, viz. Numerical Propulsion System Simulation (NPSS) that can be easily and reliably executed by an academic research group with guidance provided by a panel of experts from industry, FAA, NASA, DOE and DOD. Part 1 of this series of papers gives a summary on several topics including the need to reduce the aviation greenhouse gas (GHG) emissions, to use the alternate low carbon fuels, to develop target CO2 goal for the FAA propulsion system study and on-going technology programs providing alternatives in place of the simple Brayton cycle for contrail mitigation leading finally to the proposed advanced Brayton cycles for undertaking an extensive study under the FAA program. These studies are described in Parts 2 through 6 presented at the 10 th International Energy Conversion Engineering Conference, IECE.

An overview is given on the use of several complex multi-swirler devices in gas turbine combustion and attendant technological advances in emissions, cooling, pattern factor, operability and overall temperature increase across the combustor. The emphasis of this paper is on the initiation of promising lean-direct injection concepts, LDI and how it led to development of several innovative premix- and partially premix-concepts applicable mostly for the 30 overall pressure ratios, OPR gas turbine engines. I INTRODUCTION

Nomenclature Ce1, Ce2 – Constants in the ke equation. 2 – Jet half width t – Thermal boundary layer thickness D,E – Addition terms in the ke turbulence model e – Dissipation rate ewt Enhanced wall treatment f1, f2 – Damping functions in the „e‟ equation. h – heat transfer coefficient k – Kinetic energy l length scale P – Pressure Pk – Production of kinetic energy Re – Reynolds number RKE Realizable k- model RSM – Reynolds stress model SA – Spalart Allmaras model SKE Standard k- model SST – Shear-stress transport model swf – standard wall function Sij – Strain rate tensor T – temperature Tp – pleanum temperature Ts – surface temperature Ui – Velocity tensor

An overview is given on the use of several complex multi-swirler devices in gas turbine combustion and attendant technological advances in emissions, cooling, pattern factor, operability and overall temperature increase across the combustor. The emphasis of this paper is on the development and innovative applications of twin-concentric richand lean-direct injection mixers, their use in the dual-annular combustors, and finally making its extension to premixed mixers for the Dry Low Emissions, DLE twin and triple annular combustors.

CP – Coefficient of pressure d – diameter of the orifice ewt – Enhanced wall treatment e – Dissipation rate h – heat transfer coefficient H – distance of orifice from the target plate k – Kinetic energy L – Nozzle length Nu – Nusselt Number r – radial distance from the axis RKE Realizable k- model RSM – Reynolds stress model Re – Reynolds number Rin – orifice inlet radius Rmax – orifice outlet radius SA – Spalart Allmaras model SKE Standard k- model SST – Shear-stress transport model swf – standard wall function Ui – Velocity tensor

A comprehensive assessment has been made of the ICAO Engine Emissions Data with special emphasis on the Low-Emissions and Rich-Quench-Lean Combustors, LEC/RQL. Since the middle to late 1970’s, more than 50% reduction has been achieved in NOx, CO and smoke along with 95% reduction in the unburned hydrocarbons. All the engine original equipment manufacturers (OEM’s) through their highly inventive approaches have produced similar results as summarized in this paper. This provides an incentive for the team to define a gas turbine Relevant Combustion Simulator, RCS to reproduce gaseous and particulate emissions characteristics of LEC/RQL combustors with potential to provide guidance for developing future-gen fuel-flex, ultra-clean energy efficient engines.

A critical assessment is made of the available engineering correlations for emissions with the intention of developing the next-generation physicsbased easy to implement models for scaling combustors for the next-generation energy efficient fuel flexible aero-propulsion gas turbine engines. An engineering correlation based approach, which is considered as the 1 st step in the process of developing “applicable” physics-based simple rules and tools is described. We make use of the following four hypotheses for developing the correlations: 1. NOxEI is primarily influenced by stoichiometric flame temperature; 2. COEI is expressed as a function of NOxEI; 3. HCEI is expressed as a function of COEI for idle and low-power operation; and 4. When one is interested for the entire engine operating line, HCEI should be expressed as a function of NOxEI. The engineering correlation approach was shown to work very well for seven different combustors, and showed that for single engine emissions data of four combustors, the 2 band for NOx, CO and HC are ±4%, ±2 EI and ±0.4 EI, respectively.

This paper describes a diffuser-combustor flow interaction analysis procedure for gas turbine combustion systems. The method is based on the solution of the Navier-Stokes equations in a generalized non-orthogonal coordinate system. The turbulence effects are modeled via the standard two-equation (k-e) model. The method has been applied to a practical gas turbine combustor-diffuser system that includes support struts and fuel nozzles. Results have been presented for a three-dimensional simulation, as well as for a simplified axisymmetric simulation. The flow exhibits significant three-dimensional behavior. The axisymmetric simulation is shown to predict the static pressure recovery and the total pressure losses reasonably well.Copyright © 1990 by ASME