Joseph Zelina

Ultra-Compact Combustors for Advanced Gas Turbine Engines

Current gas turbine design practice for military and commercial aircraft is becoming marginal to meet the increasingly severe, yet conflicting requirements for reduced fuel burn, engine weight, and exhaust emissions, while achieving enhanced parts durability. The engine combustion system can be a key enabler in achieving future design goals. In both conventional single-stage and staged combustors, attention is focused on incremental fluid dynamic changes to enhance fuel-air mixing. In addition, dramatic materials development appears necessary to permit the increased operating temperatures that will ensue from advanced Brayton cycles. This paper describes a revolutionary combustion system that is far from incremental, but appears to offer the potential for continuing advances in engine performance. A gas turbine engine has been proposed that uses a near-constant-temperature (NCT) cycle and an Inter-Turbine Burner (ITB) to provide large amounts of power extraction from the low-pressure turbine. This level of energy is achieved with a modest temperature rise across the ITB. The additional energy can be used to power a large fan for an ultra-high bypass ratio transport aircraft, or to drive an alternator for large amounts of electrical power extraction. Conventional gas turbine engines cannot drive ultra-large diameter fans without the use of excessive turbine temperatures, or a substantial number of turbine stages. In addition, these conventional systems cannot meet high power extraction demands without a loss of engine thrust. The objective is to demonstrate an Ultra-Compact Combustor (UCC) that can be used as a main burner or an ITB that does not impact engine thrust-to-weight, pollutant emissions, or overall system performance. Concepts for an Ultra-Compact Combustor (UCC) are being explored experimentally. This system uses high swirl in a circumferential cavity to enhance reaction rates via high cavity g-loading on the order of 3000 g’s. Increase in reaction rates translates to a reduced combustor volume. The UCC design integrates compressor and turbine features which will enable a shorter and potentially less complex gas turbine engine. This paper will describe different variations of the UCC design where both the fuel injection method, turbine vane design, and the fuel injection angles are varied in the UCC. Experimental results from the UCC at atmospheric pressure indicate that the combustion system operates at 95–99% combustion efficiency over a wide range of operating conditions burning JP-8 +100 fuels. Axial flame lengths were extremely short, at about 50% those of conventional systems.Copyright

Emissions Reduction Technologies for Military Gas Turbine Engines

Future military gas turbine engines will have higher performance than current engines, resulting in increased compressor and combustor exit temperatures, combustor pressures, and fuel-air ratios with wider operating limits. These combustor characteristics suggest undesirable exhaust emission levels of nitrogen oxides and smoke at maximum power and higher carbon monoxide and unburned hydrocarbons at low power. To control emission levels while improving performance, durability and cost, requires major advances in combustor technology. Current emissions control approaches as applied to conventional swirl-stabilized combustors include rich- and lean-burn strategies, together with staged combustion. These approaches, even in fully developed form, may not be sufficient to satisfy the projected design requirements. Unconventional combustor configurations may become necessary. Different engine cycles other than the standard Brayton cycle may also be used for special applications in order to avoid the use of excessive combustion temperatures. The paper presents an overview of the currently utilized emissions control approaches, comparing their performances and likely potential for meeting future requirements. Experimental results are presented for two non-conventional combustor configurations that have shown promise for advanced engine applications. A brief discussion is offered on cycle changes that could result in lower peak temperatures while maintaining advanced performance.

Exploration of Compact Combustors for Reheat Cycle Aero Engine Applications

An aero gas turbine engine has been proposed that uses a near-constant-temperature (NCT) cycle and an Inter-Turbine Burner (ITB) to provide large amounts of power extraction from the low-pressure turbine. This level of energy is achieved with a modest temperature rise across the ITB. The additional energy can be used to power a large geared fan for an ultra-high bypass ratio transport aircraft, or to drive an alternator for large amounts of electrical power extraction. Conventional gas turbines engines cannot drive ultra-large diameter fans without causing excessively high turbine temperatures, and cannot meet high power extraction demands without a loss of engine thrust. Reducing the size of the combustion system is key to make use of a NCT gas turbine cycle. Ultra-compact combustor (UCC) concepts are being explored experimentally. These systems use high swirl in a circumferential cavity about the engine centerline to enhance reaction rates via high cavity g-loading on the order of 3000 g’s. Any increase in reaction rate can be exploited to reduce combustor volume. The UCC design integrates compressor and turbine features which will enable a shorter and potentially less complex gas turbine engine. This paper will present experimental data of the Ultra-Compact Combustor (UCC) performance in vitiated flow. Vitiation levels were varied from 12–20% oxygen levels to simulate exhaust from the high pressure turbine (HPT). Experimental results from the ITB at atmospheric pressure indicate that the combustion system operates at 97–99% combustion efficiency over a wide range of operating conditions burning JP-8 +100 fuel. Flame lengths were extremely short, at about 50% of those seen in conventional systems. A wide range of operation is possible with lean blowout fuel-air ratio limits at 25–50% below the value of current systems. These results are significant because the ITB only requires a small (300°F) temperature rise for optimal power extraction, leading to operation of the ITB at near-lean-blowout limits of conventional combustor designs. This data lays the foundation for the design space required for future engine designs.Copyright

Flame Stabilization in Small Cavities

This research is motivated by the necessity to improve the performance of ultracompact combustors, which requires flame stabilization in small cavities. An extensive computational investigation on the characteristics of cavity-stabilized flames is presented. A high-fidelity, time-accurate, implicit algorithm that uses a global chemical mechanism for JP8-air combustion and includes detailed thermodynamic and transport properties as well as radiation effects is used for simulation. Calculations are performed using both direct numerical simulation and standard k-e Reynolds-averaged Navier-Stokes model. The flow unsteadiness is first examined in large axisymmetric and small planar cavities with nonreactive flows. As with previous investigations on axisymmetric cavities, multiple flow regimes were obtained by varying cavity length (x/D o ) : wake backflow regime, unsteady cavity vortex regime, steady cavity vortex regime, and compressed cavity vortex regime. However, planar cavities only exhibit steady cavity vortex and compressed cavity vortex regimes. Two opposed nonaligned air jets were positioned in this planar cavity: the outermost air jet in coflow with the mainstream flow (i.e., normal injection). The fuel jet was injected either in coflow, crossflow, or counterflow with respect to the mainstream flow. Flow unsteadiness was observed to be relatively small for coflow- and crossflow-fuel-jet injection. By reversing the air jet positions (i.e., reverse injection), the flow unsteadiness is promoted regardless of fuel jet positioning. Finally, the effect of combustion and cavity equivalence ratio (φ CAV ) on flame unsteadiness is addressed. With normal injection (reverse injection), low and high φ CAV leads to low (high) and high (low) flame unsteadiness, respectively. Based on these results recommendations are provided to designers/engineers to reduce flame unsteadiness in these cavities.

Effect of Trapped Vortex Combustion with Radial Vane Cavity Arrangements on Predicted Inter-Turbine Burner Performance

The complex combustion processes, including chemical reactions, turbulence, unsteady, multiphase flow, evaporation and heat and mass transfer pose great challenges in modern propulsion system design and development. Ultra-short compact, high performance combustion systems are desirable for advanced propulsion systems from the standpoint of lower fuel consumption and increased material durability. AFRL has proposed placing an Ultra-Compact Combustor (UCC) between a high pressure turbine stage and low pressure turbine stage to create an innovative Inter-Turbine Burner (ITB) concept. This paper focuses on ITB combustor technologies that can enable the development of compact, highperformance combustion systems. Compact combustors weigh less and take up less volume in space-limited turbine engine for aero applications. The earlier designs conceived and developed at AFRL/RZTC is based on the idea that the flame speed under turbulent conditions is directly proportional to the square root of gravity and high-g flames offer increased flame speeds, which would aid in the design of shorter combustion systems. This idea led to an ITB with a circumferential cavity in which fuel and air injected at selected points led to rich combustion in the circumferential cavity. This was further followed by lean combustion and flame stabilization with the aid of a radial vane with notch. Even though this concept exhibited good merits through several rig tests and numerical studies carried out over the years at AFRL/RZTC, it does not allow scaling of the geometry and configuration for higher mass flow rates, larger size and increased thrust requirements. This paper presents an alternative concept for the UCC that uses a Trapped Vortex Cavity (TVC) to replace the high swirling circumferential cavity combustion to enhance mixing rates via a double vortex system in the TVC, followed by further mixing of the free stream air through the vane with a notch. Flow field predictions utilizing FLUENT are presented for concept evaluation in a systematic way to understand the flow development and physics, leading to the incremental combustion enhancement, total pressure loss, the entrainment and the calculated exit temperature profile. The analysis supplements the understanding of the design space required for future engine designs that may use this novel, compact combustion systems.

Operability and Efficiency Performance of Ultra-Compact, High Gravity (g) Combustor Concepts

Abstract : This paper presents a parametric design study of the Ultra-Compact Combustor (UCC), a novel design based on trapped-vortex combustor (TVC) work that uses high swirl in a circumferential cavity to enhance reaction rates via high cavity g-loading on the order of 3000 gs. Increase in reaction rates translates to a reduced combustor volume. Three combustor geometric features were varied during experiments which included (1) high-g cavity flame-holding method, (2) high-g cavity to main airflow transport method, and (3) fuel injection method. Experimental results are presented for these combustor configurations and results have shown promise for advanced engine applications. Lean blowout fuel-air ratio limits at 25-50% the value of current systems were demonstrated. Combustion efficiency was measured over a wide range of UCC operating conditions. This data begins to build the design space required for future engine designs that may use these novel, compact, high-g combustion systems.

Compact Combustion Systems Using a Combination of Trapped Vortex and High-G Combustor Technologies

Major advances in combustor technology are required to meet the conflicting challenges of improving performance, increasing durability and maintaining cost. Ultra-short combustors to minimize residence time, with special flame-holding mechanisms to cope with increased through-velocities are likely in the future. This paper focuses on vortex-stabilized combustor technologies that can enable the design of compact, high-performance combustion systems. Compact combustors weigh less and take up less volume in space-limited turbine engine for aero applications. This paper presents the UCC, a novel design based on TVC work that uses high swirl in a circumferential cavity to enhance mixing rates via high cavity g-loading on the order of 3000 g’s. The UCC design integrates compressor and turbine features which will enable a shorter and potentially less complex gas turbine engine. Ultimately, it is envisioned that this type of combustion system can be used as the main combustor and/or as a secondary combustor between the high pressure and low pressure turbine to operate as a reheat cycle engine. The focus on this paper includes experimental results of the UCC for a variety of conditions: (1) the addition of turbine vanes in the combustor flowpath, (2) a comparison of JP-8 and FT fuel performance in the combustor, (3) the use of trapped-vortex-like air addition to increase combustor flammability limits, and (4) combustor performance related to two different fuel injector designs. Lean blowout fuel-air ratio limits at 20% the value of current systems were demonstrated. Combustion efficiency was measured over a wide range of UCC operating conditions. This data begins to build the design space required for future engine designs that may use these novel, compact, high-g combustion systems.

Experimental and Computational Studies of an Ultra-Compact Combustor

Reducing the weight and decreasing pressure losses of aviation gas turbine engines improves the thrust-to-weight ratio and improves efficiency. In ultra-compact combustors (UCCs), engine length is reduced and pressure losses are decreased by merging a combustor with adjacent components using a systems engineering approach. High-pressure turbine inlet vanes can be placed in a combustor to form a UCC. Eliminating the compressor outlet guide vanes (OGVs) and maintaining swirl through the diffuser can result in further reduction in engine length and weight. Cycle analysis indicates that a 2.4% improvement in engine weight and a 0.8% increase in thrust-specific fuel consumption are possible when a UCC is used. Experiments and analysis were performed in an effort to understand key physical and chemical processes within a trapped-vortex UCC. Experiments were performed using a combustor operating at pressures in the range of 520–1030 kPa (75–150 psi) and inlet temperature of 480–620 K (865–1120 °R). The primary reaction zone is in a single trapped-vortex cavity where the equivalence ratio was varied from 0.7 to 1.8. Combustion efficiencies and NOx emissions were measured and exit temperature profiles obtained, for various air loadings, cavity equivalence ratios, and configurations with and without turbine inlet vanes. A combined diffuser-flameholder (CDF) was used in configurations without vanes to study the interaction of cavity and core flows. Higher combustion efficiency was achieved when the forward-to-aft momentum ratios of the air jets in the cavity were near unity or higher. Discrete jets of air immediately above the cavity result in the highest combustion efficiency. The air jets reinforce the vortex structure within the cavity, as confirmed through coherent structure velocimetry of high-speed images. A more uniform temperature profile was observed at the combustor exit when a CDF is used instead of vanes. This is the result of increased mass transport along the face of the flame holder. Emission indices of NOx were between 3.5 and 6.5 g/kgfuel for all test conditions. Ultra-compact combustors (with a single cavity) can be run with higher air loadings than those employed in previous testing with a trapped-vortex combustor (two cavities) with similar combustion efficiencies being maintained. The results of this study suggest that the length of combustors and adjacent components can be reduced by employing a systems level approach.Copyright

High -Pressure Tests of a High -g, Ultra -Compact Combustor

The Ultra-Compact Combustor (UCC) is part of evolving technology in the development of near-constant-temperature-cycle gas turbine engines. This technology can provide a significant reduction in engine weight and size while providing large amounts of power. The UCC uses high swirl in a circumferential cavity to enhance reaction rates via high cavity gloading on the order of 3000 g’s. Increase in reaction rates translates to a reduced combustor volume. Axial flame lengths are extremely short, at about 50% those of conventional systems. High-pressure UCC tests conducted in the Air Force Research Laboratory (AFRL) High Pressure Combustor Research Facility (HPCRF) have demonstrated the feasibility of using UCC technology in advanced main combustor and Inter-Turbine Burner (ITB) systems. The UCC design integrates compressor and turbine features which will enable a shorter and less complex gas turbine engine. Experimental results from UCC testing at elevated pressure indicated that the combustion system operates at 95-99 percent efficiency over an increased operating range compared to conventional gas turbine combustion systems burning JP-8+100 fuels. This paper will describe experimental results from one UCC configuration.

The Impact of Heat Release in Turbine Film Cooling

The Ultra Compact Combustor is a design that integrates a turbine vane into the combustor flow path. Because of the high fuel-to-air ratio and short combustor flow path, a significant potential exists for unburned fuel to enter the turbine. Using contemporary turbine cooling vane designs, the injection of oxygen-rich turbine cooling air into a combustor flow containing unburned fuel could result in heat release in the turbine and a large decrease in cooling effectiveness. The current study explores the interaction of cooling flow from typical cooling holes with the exhaust of a fuel-rich well-stirred-reactor operating at high temperatures over a flat plate. Surface temperatures, heat flux, and heat transfer coefficients are calculated for a variety of reactor fuel-to-air ratios, cooling hole geometries, and blowing ratios. Results demonstrate that reactions in the turbine cooling film can result in increased heat transfer to the surface. The amount of this increase depends on hole geometry and blowing ratio and fuel content of the combustor flow. Failure to design for this effect could result in augmented heat transfer caused by the cooling scheme, and turbine life could be degraded substantially.