Combustor-turbine Interface Research Articles

Investigation of Unsteady Flow Phenomena in First Vane Caused by Combustor Flow With Swirl

The flow at the combustor turbine interface of power generation gas turbines with can combustors is characterized by high and nonuniform turbulence levels, lengthscales, and residual swirl. These complexities have a significant impact on the first vanes aerothermal performance and lead to challenges for an effective turbine design. To date, this design philosophy mostly assumed steady flow and thus largely disregards the intrinsic unsteadiness. This paper investigates the steady and unsteady effects of the combustor flow with swirl on the turbines first vanes. Experimental measurements are conducted on a high-speed linear cascade that comprises two can combustors and four nozzle guide vanes (NGVs). The experimental results are supported by a large eddy simulation (LES) performed with the inhouse computational fluid dynamics (CFD) flow solver TBLOCK. The study reveals the highly unsteady nature of the flow in the first vane and its effect on the heat transfer. A persistent flow structure of concentrated vorticity is observed. It wraps around the unshielded vane's leading edge (LE) at midspan and periodically oscillates in spanwise direction due to the interaction of the residual low-pressure swirl core and the vane's potential field. Moreover, the transient behavior of the horseshoe-vortex system due to large fluctuations in incidence is demonstrated.

Heat transfer measurement near endwall region of first stage gas turbine nozzle having platform misalignment at combustor-turbine interface

The effect of a misalignment between vane endwall and combustor exit in a gas turbine was investigated using a Computational Fluid Dynamics (CFD) simulation and experimental measurements. The misaligned endwall platform was simulated as a backward facing step in this study. The CFD simulation predicted two legs of the vortex, referred to as a step-induced vortex, created by the step flowing through nozzle passage. Heat transfer measurements demonstrated the effect of the step-induced vortex on the endwall and the vane surface indicated by locally increased heat transfer coefficients which corresponded to the locus of the vortex, as also predicted by the simulation. Although a boundary layer transition occurred early, the locally increased heat transfer persisted to the vane trailing edge. In summary, a misaligned endwall platform causes negative effects on the gas turbine with respect to the thermal design. A vortex was generated by the step, which caused a higher thermal load on the nozzle vane surfaces, especially near the endwall.

Advanced Combustor Exit Plane Temperature Diagnostics Based on Large Eddy Simulations

The accurate description and flow calibration at the exit of a gas turbine combustor is a key point of turbine design. Indeed, the temperature field at the turbine inlet directly impacts both the work extraction by the turbine stages as well as drives the wall heat transfer occurring on the vanes and rotor surfaces. Properly simulating or measuring this critical planar interface has been an extensive topic of research. Despite such efforts, industrial exploitation of the various measures at that interface usually resumes to a quantification of the temperature heterogeneities through the Local Radial Temperature Distortion Factor (LRTDF), thereby loosing major spatial and unsteady information. As lean burn configurations become a standard for aero-engine combustors, the absence of dilution holes reduces the intensity of mixing and the use of more advanced statistical tools should propose an exhaustive representation of the temperature map and mixing quality. Current state-of-the-art numerical design of gas turbine combustor relies on Large Eddy Simulations (LES) giving access to a fully 3-D temporally dependent flow field throughout the combustion chamber. Going beyond LRTDF is therefore possible and advanced analyses of the combustor-turbine interface are accessible to better qualify proposed design changes. In this work, a simple statistical representation of the temperature distribution is introduced and applied for the analysis of a lean burn combustor simulator interface. It is shown that at the exit of this chamber, the temperature samples are distributed following a non-Gaussian shape, highlighting a most probable local temperature largely different from the conventional mean value. The use of high order moments explains such deviations and indicates the presence of strongly segregated zones governed by the technological implementation of effusion cooling systems. Based on this diagnostic, a new and complementary representation of the LRTDF is proposed, accounting for the statistical dimension of the classical (radius, temperature) formulation.

Impact of the Combustor-Turbine Interface Slot Orientation on the Durability of a Nozzle Guide Vane Endwall

The combustor-turbine interface is an essential component in a gas turbine engine as it allows for thermal expansion between the first stage turbine vanes and combustor section. Although not considered as part of the external cooling scheme, leakage flow from the combustor-turbine interface can be utilized as coolant. This paper reports on the effects of orientation of a two-dimensional leakage slot, simulating the combustor-turbine interface, on the net heat flux reduction to a nozzle guide vane endwall. In addition to adiabatic effectiveness and heat transfer measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed in the vane stagnation plane. Four interface slot orientations of 90 deg, 65 deg, 45 deg, and 30 deg located at 17% axial chord upstream of a first vane in a linear cascade were studied. Results indicate that reducing the slot angle to 45 deg can provide as much as a 137% reduction to the average heat load experienced by the endwall. Velocity measurements indicate the formation of a large leading edge vortex for coolant injected at 90 deg and 65 deg while coolant injected at 45 deg and 30 deg flows along the endwall and washes up the vane surface at the endwall junction.

Turbine endwall film cooling with combustor-turbine interface gap leakage flow: Effect of incidence angle

This paper is focused on the film cooling performance of combustor-turbine leakage flow at off-design condition. The influence of incidence angle on film cooling effectiveness on first-stage vane endwall with combustor-turbine interface slot is studied. A baseline slot configuration is tested in a low speed four-blade cascade comprising a large-scale model of the GE-E 3 Nozzle Guide Vane (NGV). The slot has a forward expansion angle of 30 deg. to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5 × 10 5 and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet). The blowing ratio of the coolant through the interface gap varies from M = 0.1 to M = 0.3, while the blowing ratio varies from M = 0.7 to M = 1.3 for the endwall film cooling holes. The film-cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique. The results show that with an increasing blowing ratio the film-cooling effectiveness increases on the endwall. As the incidence angle varies from i = +10 deg. to i = 10 deg., at low blowing ratio, the averaged film-cooling effectiveness changes slightly near the leading edge suction side area. The case of i = +10 deg. has better film-cooling performance at the downstream part of this region where the axial chord is between 0.15 and 0.25. However, the disadvantage of positive incidence appears when the blowing ratio increases, especially at the upstream part of near suction side region where the axial chord is between 0 and 0.15. On the main passage endwall surface, as the incidence angle changes from i = +10 deg. to i = 10 deg., the averaged film-cooling effectiveness changes slightly and the negative incidence appears to be more effective for the downstream part film cooling of the endwall surface where the axial chord is between 0.6 and 0.8.

Effects of Orientation and Position of the Combustor-Turbine Interface on the Cooling of a Vane Endwall

First stage, nozzle guide vanes and accompanying endwalls are extensively cooled by the use of film cooling through discrete holes and leakage flow from the combustor-turbine interface gap. While there are cooling benefits from the interface gap, it is generally not considered as part of the cooling scheme. This paper reports on the effects of the position and orientation of a two-dimensional slot on the cooling performance of a nozzle guide vane endwall. In addition to surface thermal measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed at the vane stagnation plane. Two slot orientations, 90 deg and 45 deg, and three streamwise positions were studied. Effectiveness results indicate a significant increase in area averaged effectiveness for the 45 deg slot relative to the 90 deg orientation. Flowfield measurements show dramatic differences in the horseshoe vortex formation.

Experimental and Numerical Investigation of Combustor-Turbine Interaction Using an Isothermal, Nonreacting Tracer

Understanding the interaction between the combustor and turbine subsystems of a gas turbine engine is believed to be key in developing focused strategies for improving turbine performance. Past studies have approached the problem starting with an existing turbine rig with inlet conditions provided by “representative” hardware which attempts to mimic some key flow features exiting the combustor. In this paper, experiments are performed which center around complete engine hardware of the combustor, including engine geometry turbine nozzle guide vanes (NGVs) to solely represent the upstream impact of the complete turbine. This domain ensures that the traditional interface between combustor and turbine is sufficiently encompassed and not compromised by obfuscating or limiting effects due to approximating combustor hardware. The full-annular experimental measurements include all components of the velocity and pressure fields at various planar sections perpendicular to the primary flow direction. These include detailed, two-dimensional measurements both upstream and downstream of the NGVs. The combustor is a classic rich-burn design. Passive scalar (CO2) tracing measurements are performed to gain insight into the flow responsible for the temperature fields in the coupled system, including the impact of the NGVs on the upstream flow at the conventional combustor-turbine interface. CFD simulations are used to develop a complete picture of the combustor-turbine interface and the coupling between the two subsystems. The complementary experimental and simulation datasets are together intended to provide a benchmark for future, more traditional turbine rig tests and turbine CFD simulations where inlet conditions are at the exit plane of the combustor.

The Effect of the Combustor-Turbine Slot and Midpassage Gap on Vane Endwall Heat Transfer

Turbine vanes are generally manufactured as single- or double-airfoil sections that are assembled into a full turbine disk. The gaps between the individual sections, as well as a gap between the turbine disk and the combustor upstream, provide leakage paths for relatively higher-pressure coolant flows. This leakage is intended to prevent ingestion of the hot combustion flow in the primary gas path. At the vane endwall, this leakage flow can interfere with the complex vortical flow present there and thus affect the heat transfer to that surface. To determine the effect of leakage flow through the gaps, heat transfer coefficients were measured along a first-stage vane endwall and inside the midpassage gap for a large-scale cascade with a simulated combustor-turbine interface slot and a midpassage gap. For increasing combustor-turbine leakage flows, endwall surface heat transfer coefficients showed a slight increase in heat transfer. The presence of the midpassage gap, however, resulted in high heat transfer near the passage throat where flow is ejected from that gap. Computational simulations indicated that a small vortex created at the gap flow ejection location contributed to the high heat transfer. The measured differences in heat transfer for the various midpassage gap flowrates tested did not appear to have a significant effect.

Film-Cooling Flowfields With Trenched Holes on an Endwall

The leading edge region along the endwall of a stator vane experiences high heat transfer rates resulting from the formation of horseshoe vortices. Typical gas turbine endwall designs include a leakage slot at the combustor-turbine interface as well as film-cooling holes. Past studies have documented the formation of a horseshoe vortex at the leading edge of a vane, but few studies have documented the flowfield in the presence of an interface slot and film-cooling jets. In this paper, a series of flowfield measurements is evaluated at the leading edge with configurations including a baseline with neither film-cooling holes nor an upstream slot, a row of film-cooling holes and an interface slot, and a row of film-cooling holes in a trench and an interface slot. The results indicated the formation of a second vortex present for the case with film-cooling holes and a slot relative to the baseline study. In addition, turbulence intensity levels as high as 50% were measured at the leading edge with film-cooling holes and a slot compared with the 30% measured for the baseline study. A trench was shown to provide improved overall cooling relative to the no trench configuration as more of the coolant stayed attached to the endwall surface with the trench.

The Effects of Varying the Combustor-Turbine Gap

To protect hot turbine components, cooler air is bled from the high pressure section of the compressor and routed around the combustor where it is then injected through the turbine surfaces. Some of this high pressure air also leaks through the mating gaps formed between assembled turbine components where these components experience expansions and contractions as the turbine goes through operational cycles. This study presents endwall adiabatic effectiveness levels measured using a scaled up, two-passage turbine vane cascade. The focus of this study is evaluating the effects of thermal expansion and contraction for the combustor-turbine interface. Increasing the mass flow rate for the slot leakage between the combustor and turbine showed increased local adiabatic effectiveness levels while increasing the momentum flux ratio for the slot leakage dictated the coverage area for the cooling. With the mass flow held constant, decreasing the combustor-turbine interface width caused an increase in uniformity of coolant exiting the slot, particularly across the pressure side endwall surface. Increasing the width of the interface had the opposite effect thereby reducing coolant coverage on the endwall surface.

Effect of Midpassage Gap, Endwall Misalignment, and Roughness on Endwall Film-Cooling

To maintain acceptable turbine airfoil temperatures, film cooling is typically used whereby coolant, extracted from the compressor, is injected through component surfaces. In manufacturing a turbine, the first stage vanes are cast in either single airfoils or double airfoils. As the engine is assembled, these singlets or doublets are placed in a turbine disk in which there are inherent gaps between the airfoils. The turbine is designed to allow outflow of high-pressure coolant rather than hot gas ingestion. Moreover, it is quite possible that the singlets or doublets become misaligned during engine operation. It has also become of interest to the turbine community as to the effect of corrosion and deposition of particles on component heat transfer. This study uses a large-scale turbine vane in which the following two effects are investigated: the effect of a midpassage gap on endwall film cooling and the effect of roughness on endwall film cooling. The results indicate that the midpassage gap was found to have a significant effect on the coolant exiting from the combustor-turbine interface slot. When the gap is misaligned, the results indicate a severe reduction in the film-cooling effectiveness in the case where the pressure side endwall is below the endwall associated with the suction side of the adjacent vane.