Performance improvements of conventional gas turbines are becoming increasingly difficult and costly to achieve. Pressure Gain Combustion (PGC) has emerged as a promising technology in this respect, due to the higher thermal efficiency of the respective ideal gas turbine cycle. Previous cycle analyses considering turbine cooling methods have shown that the application of pressure gain combustion may require more turbine cooling air. This has a direct impact on the cycle efficiency and reduces the possible efficiency gain that can potentially be harvested from the new combustion technology. Novel cooling techniques could unlock an existing potential for a further increase in efficiency. Such a novel turbine cooling approach is the application of pulsed impingement jets inside the turbine blades. In the first part of this paper, results of pulsed impingement cooling experiments on a curved plate are presented. The potential of this novel cooling approach to increase the convective heat transfer in the inner side of turbine blades is quantified. The second part of this paper presents a gas turbine cycle analysis where the improved cooling approach is incorporated in the cooling air calculation. The effect of pulsed impingement cooling on the overall cycle efficiency is shown for both Joule and PGC cycles. In contrast to the authors’ anticipation, the results suggest that for relevant thermodynamic cycles pulsed impingement cooling increases the thermal efficiency of Joule cycles more significantly than it does in the case of PGC cycles. Thermal efficiency improvements of 1.0 p.p. for pure convective cooling and 0.5 p.p. for combined convective and film with TBC are observed for Joule cycles. But just up to 0.5 p.p. for pure convective cooling and 0.3 p.p. for combined convective and film cooling with TBC are recorded for PGC cycles.

This paper numerically investigated the effect of the arrays of the delta-winglet vortex generators (DWVGs) pairs on the flow field and heat transfer characteristics of gas turbine blade tip internal surface. Six different arrangements including three inclinations (30°, 45°, 60°) and two aspect ratios of DWVGs are calculated at Reynolds numbers ranging from 6000 to 14,000. The internal cooling passage of gas turbines are simplified as two-pass U bend channel and the U bend channel without any tabulators are considered as Baseline. The detailed flow structure, the evolution of vortices and heat transfer performance over the tip internal surface are presented. The results show that the arrays of DWVGs placed on the tip internal surface have great influence on the tip flow and heat transfer. Small-scale vortices are induced by the DWVGs, which have negligible impact on the main flow. Due to the nature of 180-deg turn, the impingement-like flow contributes the highest heat transfer performance. But too many DWVGs placed on the attachment region will weaken the energy of main flow and therefore reduce the local heat transfer. Besides, the blocked DWVGs (BVG) will enhance the heat transfer at the center line, and the guided DWVGs (GVG) will extend the low-energy flow cluster and thus weakening the heat transfer intensity. The results of this study are useful for understanding the mechanism of heat transfer characteristics in a realistic gas turbine blade by using the arrays of DWVGs.

High aspect ratio channels are a common internal cooling feature in Gas Turbine blades, mostly suitable for the trailing edge region or mid-chord regions. Traditionally such channels are fitted with rib-turbulators and/or pin-fin turbulators to augment heat transfer and prevent material failure. Highly efficient internal cooling of blades can improve the efficiency of a real Gas Turbine power cycle by tolerating higher Turbine Inlet Temperatures (TIT). Multi-physics Topology optimization (TO) has been employed in the current study to find optimized shape of these ducts, with an aim to increase heat transfer, while constraining the pressure drop across the channel. This method, commonly used in structural problems, is a novel topic of research when applied to fluid-thermal studies. Material distribution in the computational domain is varied by changing porosity value in each cell and thereby altering the fluid path and creating a conjugate heat transfer problem. Each cell has a value of Brinkmann porosity factor which either simulates a blockage, or a fluid region depending on a low or high value of this design variable. Hence the degree of freedom is high in this method, and there is no manual bias introduced, unlike in parametric shape optimization which is limited to a few design parameters. The unconventional geometries obtained as an end product of this optimization process can thus be an alternative to existing rib/pin-fin type of cooling geometries. The recent progress in additive manufacturing can now facilitate the manufacturing of complicated shapes. An in-house Open-FOAM solver has been used to carry out the process in only twice the amount of time compared to a regular RANS-CFD. 3-Dimensional rectangular channels with inlet aspect ratios of 4:1 and 8:1 have been considered as baselines with a constant inlet velocity. Resulting optimum geometries were found to have organic tree like branching arrangements of rib-like wall roughness and v-shaped structures.

Internal cooling channels with pin-fin arrays are an important part of gas turbine blade trailing edge design. Short pin-fins act as turbulators in high aspect ratio channels to increase heat transfer and provide structural support to the trailing edge of the blade. Such pin fins however also result in a high pressure drop owing to chaotic flow phenomenon in these highly turbulent flows. High pressure-drop results in higher compressor work due to increased power consumption to push the coolant through these passages. Hence, optimizing the design of pin fin arrays is key to increasing the efficiency of real gas turbine cycles by handling higher turbine inlet temperature and increasing blade life. Moreover, the design process of such pin fin arrays can be computationally very expensive, since it typically involves high-fidelity CFD simulations. The optimization problem involves maximizing Nusselt number, while keeping the friction factor as a constraint. To address this problem, a computationally efficient approach involving Gaussian Processes (GP) surrogate modeling and constrained Bayesian Optimization (BO) has been carried out for optimizing the thermal performance of the pin fin arrays. The multidimensional search space of design parameters includes pin-fin dimensions and shape of the resulting pin-fins. The optimization problem is solved under computational budget limitations and design constraints. A ‘drop’ like optimal design is obtained which has a lower pressure drop and higher Nu compared to the baseline.

This work reports the results of paired experiments for a complex internal cooling flow within a gas turbine vane using Magnetic Resonance Velocimetry (MRV) and steady-state Infrared (IR) thermometry. A scaled model of the leading edge insert for a gas turbine vane with multi-pass impingement was designed, built using stereolithography (SLA) fabrication methods, and tested using MRV techniques to collect a three-dimensional, three-component velocity field data set for a fully turbulent test case. Stagnation and recirculation zones were identified and assessed in terms of impact on potential cooling performance. A paired experiment employed an IR camera to measure the temperature profile data of a thin, heated stainless steel impingement surface modeling the inside turbine blade wall cooled by the impingement from the vane cooling insert, providing complementary data sets. The temperature data allow for the calculation of wall heat transfer characteristics, including the Nusselt number distribution for cooling performance analysis to inform design and validate computational models. Quantitative and qualitative comparisons of the paired results show that the flow velocity and cooling performance are highly coupled. Module-to-module variation in the surface Nusselt number distributions are evident, attributable to the complex interaction between transverse and impinging flows within the apparatus. Finally, a comparison with internal heat transfer correlations is conducted using the data from Florschuetz [1]. Measurement uncertainty was assessed and estimated to be approximately +7% for velocity and ranging from +3% to +10% for Nusselt number.

In gas turbines, the turbine blades are always working in the highly temperature overhead the permissible metal temperatures. To safe operation, the turbine blades are needed to cool. Many researchs in turbine cooling technology can be categorized as internal and external cooling. This paper presents an investigation of cutted-root rib design, where a part of rib was truncated below to create an extra-passage in the root rib applied in the internal cooling turbine blades of jet engine using three-dimensional Reynolds-averaged Navier-Stokes with the SST model. The object of this investigation is to reduce the vortex occurring near the rib for improving the performance of heat transfer, such as the Nusselt number and thermal performance factor. To investigate the heat transfer performance and fluid flow characteristics of internal cooling turbine blades, a parametric study of the cutted-root rib was performed using various geometric parameters related to the height and shapes of the extra-passage. The cutted-root rib geometry is designed in ANSYS DesignModeler, and then meshed by using ICEM-CFD, analysed and post-processed using Ansys-CFX. The numerical results showed that all heat transfer parameter with the cutted-root rib design was greater than the original case without cutted-root rib.

It is well-known from the literature that surface roughness affects significantly friction and heat transfer. This is even more evident for additive manufactured (AM) components, which are taking an increasingly important role in the gas turbine field. However, the exploitation of numerical approaches to improve their design is hindered by the lack of dedicated correlations and CFD model developed for such high roughness conditions. Usually the additive manufactured components are simulated considering the surfaces as smooth or applying an equivalent sand-grain roughness (k s ) that results in a velocity shift in the boundary layer. However, determining a priori the most appropriate value of k s is challenging, as dozens of correlations are available, returning scattered and uncertain results. The aim of this work is to benchmark some existing modelling strategies (among which the equivalent sand grain roughness) and test a numerical approach capable of narrowing the existing gap between simulated and tested thermal performance of additive manufactured devices. The technology enabler is represented by higher-fidelity CFD simulations accounting for the impact of real surface roughness on pressure drop and heat transfer. At this purpose, an existing literature model for rough walls has been implemented in ANSYS Fluent and tested on a variety of AM mini-channels so as to determine the best-fitting values of k s and corrected wetted surface ratio S corr that match the experimental data in terms of friction factor and Nusselt number. Knowing also the measured roughness descriptors of each component, it has been possible to derive valuable guidelines for an effective exploitation of CFD on additive manufactured components, thus allowing a more accurate estimation of the thermal performance in additive manufactured components.

As designers aim to increase efficiency in gas turbines for aircraft propulsion and power generation, spatially-resolved experimental measurements are needed to validate computational models and compare improvement gains of new cooling designs. Infrared (IR) thermography is one such method for obtaining spatially-resolved temperature measurements. As technological advances in thermal detectors enable faster integration times, surface temperature measurements of rotating turbine blades become possible to capture including the smallest features. This paper outlines opportunities enabled by the latest IR detector technologies for capturing spatially-resolved rotating blade temperatures, while also addressing some of the challenges of implementing IR for turbine rigs such as the one in the Steady Thermal Aero Research Turbine (START) Laboratory. This paper documents critical steps in achieving accurate measurements including calibration, integration times, spatial noise, and motion blur. From these results, recommendations are provided for achieving accurate IR measurements collected in a rotating turbine facility to study film cooling.

Multi-row film cooling is widely applied in modern gas turbine cooling designs. However, there are few effective methods to predict local cooling effectiveness and the Sellers model based on the existing 1-D correlation is popular in the community but its prediction accuracy is questionable in some cases. The primary work of the current study is to analyze the prediction error of the Sellers model and to reveal the underlying flow mechanism. Considering the lateral uniformity assumption of the 1-D Sellers model, the effect of pitchwise row-to-row misalignment can hardly be included and the 2-D Sellers model is, therefore, proposed. Numerical simulations under density ratio near 1.0 are conducted on the double-row film cooling. Configurations of inline and staggered arrangements and blowing ratios from 0.5 to 2.0 are included. Cooling effectiveness can be theoretically decomposed into two parts contributed by separate rows, and this method has been validated. The footprint of coolant can, therefore, be tracked. Based on vortical interaction analysis, conclusions are obtained that pitchwise row-to-row spacing determines the overestimation or underestimation of the model-predicted result and the blowing ratio affects the absolute value of the prediction error. By comparing prediction accuracy of the two rows, the conclusion can be drawn that the prediction error is, to a greater extent, contributed by the row #2 because it is more likely to be affected by the row-to-row interaction.

Ceramic matrix composites (CMCs) are quickly becoming more prevalent in the design of gas turbines due to their advantageous weight and thermal properties. While there are many advantages, the CMC surface morphology differs from that of conventional cast airfoil components. Despite a great deal of research focused on the material properties of CMCs, little public work has been done to investigate the impact that the CMC surface morphology has on the boundary layer development and resulting heat transfer. In this study, a scaled-up CMC weave pattern was developed and tested in a low speed wind tunnel to evaluate both heat transfer and boundary layer characteristics. Results from these experiments indicate that the CMC weave pattern results in augmented heat transfer and flow field properties that significantly vary locally when compared to a smooth surface.

To reduce the emission of greenhouse gases, especially carbon dioxide, the combustion of hydrogen and oxy-fuel is a development direction for the gas turbine systems. Because the fuels are different from the traditional gas turbine which uses natural gas as fuel, the compositions of the working fluid are different, which can lead to differences in heat transfer characteristics. Furth more, in order to get higher overall efficacy, the turbine inlet temperature is becoming higher and higher. The radiation becomes an important factor then must be considered seriously. To study the influence of different gas compositions on heat transfer characteristics, three gas compositions, representing the combustion products of IGCC system, hydrogen turbine system, and the oxy-fuel turbine system, respectively, is chosen in this paper. Based on a real turbine first stage vane, the influence of radiation and gas compositions are both researched using CFD simulations. The results show that different gas compositions can lead to different heat transfer characteristics on the vane. The radiation effect can be dealt with as an independent factor in the turbine heat transfer analysis.

This paper presents the development and implementation of a new generation of double-sided heat-flux gauges at The Ohio State University Gas Turbine Laboratory (GTL) along with heat transfer measurements for film-cooled airfoils in a single-stage high-pressure transonic turbine operating at design corrected conditions. Double-sided heat flux gauges are a critical part of turbine cooling studies, and the new generation improves upon the durability and stability of previous designs while also introducing high-density layouts that provide better spatial resolution. These new customizable high-density double-sided heat flux gauges allow for multiple heat transfer measurements in a small geometric area such as immediately downstream of a row of cooling holes on an airfoil. Two high-density designs are utilized: Type A consists of 9 gauges laid out within a 5 mm by 2.6 mm (0.20 inch by 0.10 inch) area on the pressure surface of an airfoil, and Type B consists of 7 gauges located at points of predicted interest on the suction surface. Both individual and high-density heat flux gauges are installed on the blades of a transonic turbine experiment for the second build of the High-Pressure Turbine Innovative Cooling program (HPTIC2). Run in a short duration facility, the single-stage high-pressure turbine operated at design-corrected conditions (matching corrected speed, flow function, and pressure ratio) with forward and aft purge flow and film-cooled blades. Gauges are placed at repeated locations across different cooling schemes in a rainbow rotor configuration. Airfoil film-cooling schemes include round, fan, and advanced shaped cooling holes in addition to uncooled airfoils. Both the pressure and suction surfaces of the airfoils are instrumented at multiple wetted distance locations and percent spans from roughly 10% to 90%. Results from these tests are presented as both time-average values and time-accurate ensemble averages in order to capture unsteady motion and heat transfer distribution created by strong secondary flows and cooling flows.

High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

In gas turbine engines, the first-stage vanes usually suffer harsh incoming flow conditions from the combustor with high pressure, high temperature and high turbulence. The combustor-generated high freestream turbulence and strong secondary flows in a gas turbine vane passage have been reported to augment the endwall thermal load significantly. This paper presents a detailed numerical study on the effects of high freestream turbulence intensity, turbulence length scale, and exit Reynolds number on the endwall secondary flow pattern and heat transfer distribution of a transonic linear turbine vane passage at realistic engine Mach numbers, with a flat endwall no cooling. Numerical simulations were conducted at a range of different operation conditions: six freestream turbulence intensities (Tu = 1%, 5%, 10%, 13%, 16% and 20%), six turbulence length scales (normalized by the vane pitch of Λ/P = 0.01, 0.04, 0.07, 0.12, 0.24, 0.36), and three exit isentropic Mach number (Ma ex = 0.6, 0.85 and 1.02 corresponding exit Reynolds number Re ex = 1.1 × 10 6 , 1.7 × 10 6 and 2.2 × 10 6 , respectively, based on the vane chord). Detailed comparisons were presented for endwall heat transfer coefficient distribution, endwall secondary flow field at different operation conditions, while paying special attention to the link between endwall thermal load patterns and the secondary flow structures. Results show that the freestream turbulence intensity and length scale have a significant influence on the endwall secondary flow field, but the influence of the exit Reynolds number is very weak. The Nusselt number patterns for the higher turbulence intensities (Tu = 16%, 20%) appear to be less affected by the endwall secondary flows than the lower turbulence cases. The thermal load distribution in the arc region around the vane leading edge and the banded region along the vane pressure side are influenced most strongly by the freestream turbulence intensity. In general, the higher freestream turbulence intensities make the vane endwall thermal load more uniform. The Nusselt number distribution is only weakly affected by the turbulence length scale when Λ/P is larger than 0.04. The heat transfer level appears to have a significant uniform augmentation over the whole endwall region with the increasing Ma ex . The endwall thermal load distribution is classified into four typical regions, and the effects of freestream turbulence, exit Reynolds number in each region were discussed in detail.

This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

In the present study, a compact heat exchanger for cryogenically fueled gas turbine engine applications is introduced. The proposed concept can be integrated into one or various vanes that comprise the compression system and uses the existing vane surface to reject core heat to the cryogenic fuel. The requirements for the heat exchanger are defined for a large geared-turbofan engine operating on liquid hydrogen. The resulting preliminary conceptual design is integrated into a modified interconnecting duct and connected to the last stage of a publicly available low-pressure compressor geometry. The feasibility of different designs is investigated numerically, providing a first insight on the parameters that govern the design of such a component.

Gas turbine blades and vanes are typically manufactured with small clearances between adjacent vane and blade platforms, termed the midpassage gap. The midpassage gap reduces turbine efficiency and causes additional heat load into the vane platform, as well as changing the distribution of endwall heat transfer and film cooling. This paper presents a low-order analytical analysis to quantify the effects of the midpassage gap on aerodynamics and heat transfer, verified against an experimental campaign and CFD. Using this model, the effects of the gap can be quantified, for a generic turbine stage, based only on geometric features and the passage static pressure field. It is found that at present there are significant losses and a large proportion of heat load caused by the gap, but that with modified design this could be reduced to negligible levels. Cooling flows into the gap to prevent ingression are investigated analytically and with CFD. Recommendations are given for targets that turbine designers should work toward in reducing the adverse effects of the midpassage gap. A method to estimate the effect of gap flow is presented, so that for any machine the significance of the gap may be assessed.

As a civil gas turbine cools down, asymmetric natural convective heat transfer causes the bottom sector of the rotor to cool faster than the top; this circumferential thermal gradient can potentially cause the shaft to deflect — a phenomenon called thermal or rotor bow. Rotor bow is tremendously difficult to predict due to its dependence on a number of engine design parameters, in addition to the complex nature of natural convective flows. A novel experimental facility has been developed to gain further understanding into shutdown cooling of a gas turbine. The scope of this paper is to quantify the effect of basic design features on natural convective cooling in an engine annulus during shut-down; specifically the influence of the thermal boundary wall conditions and the annular diameter ratio. In addition to this, a low-cost, robust thermocouple probe has been developed and validated, which allows for accurate temperature measurements in a natural convective boundary layer. An extensive experimental campaign has been completed. The key finding is that the local radial wall temperature difference was found to be the most influential parameter on the local heat transfer. Non-isothermal walls did not alter the overall distribution of the inner wall equivalent conductivity. This was true for both cylindrical and conical sections. An appropriate characteristic length for use in the Rayleigh number definition for both the concentric cylinder and conical sections have been validated. The conical inner section (5° hade angle) did not affect the overall heat transfer in the range of conditions tested. Therefore, the mean surface heat transfer for non-isothermal inner and outer profiles, within the range −0.4 < ΔRa/Ra L c < 0.4, where the thermal gradient is negative in the clockwise from top-dead-centre, can be predicted using isothermal correlations for Ra L c < 5.0 × 10 5 and Dr <= 1.5.

As gas turbine inlet temperature continuously increases, blade trailing edge suffers an extremely high thermal load due to the thin structure and constraint of internal convective cooling arrangements. To overcome these difficulties, pressure-side cutback, which is strengthened with multiple internal structures and land extensions, is widely used in blades to protect trailing edge from high thermal stresses. Due to the geometrical complexity and strong interactions between coolant flow and mainstream, sophisticated heat transfer and flow patterns exist in the cutback region, which presents a great challenge for the trailing edge cutback design. To understand the heat transfer and aerodynamic performance in blade with trailing edge cutback, CFD method has become an efficient tool which provides deep insights into the flow mechanisms and heat transfer characteristics in the detailed region. To accurately resolve the flow and heat transfer performance in a turbine blade with trailing edge cutback, structured grids are preferred because of higher resolution in flow/heat transfer prediction than unstructured grids, especially in boundary layers. However, for a blade with landed trailing edge cutback, few researchers tried to employ structured grids to predict aero-thermal performance due to the geometrical complexity. In this paper, the Background-Grid Based Mapping (BGBM) method proposed in Part I of this study was adopted to generate multi-block structured grids for a gas turbine vane with landed trailing edge cutback. With the coordinate transformation strategies, multi-block structured grids for the vane with landed trailing edge cutback were generated conveniently. With the generated structured grids, flow and heat transfer performance in vane were investigated using RANS (Reynolds-Averaged Navier-Stokes) equations solutions combined with transitional turbulence model. Effects of land extensions on the heat transfer and aerodynamic performance were analyzed, as well as the effects of inflow turbulence intensity, mainstream Reynolds number and ejection rate. The results show that heat transfer coefficients on vane surface, total pressure loss coefficient and energy loss coefficient in vane are all increased with the increase of inflow turbulence intensity. However, heat transfer coefficients on cutback and trailing edge surface are not sensitive to inflow turbulence intensity. At the same inflow turbulence intensity, the aerodynamic loss in vane is decreased with increasing the Reynolds number of mainstream. The increase of ejection rate significantly increases the heat transfer coefficients on cutback surface. Compared with the vane without land extensions, heat transfer coefficients and pressure coefficients on vane surface are reduced and the heat transfer coefficients on cutback surface are increased for the vane with land extensions. In the case of Re = 2.0 × 10 6 , the area-averaged heat transfer coefficient on landed cutback is 14.46% higher than the cutback without lands. Compared with the experimental data, predictions with structured grids based on BGBM method are more agreeable than those with unstructured grids.

The effects of sinusoidal oscillations in the main flow on film cooling in the gas turbine were investigated by Large Eddy Simulation (LES). The film cooling flow fields for the sinusoidal oscillation of 32 Hz in the mainstream from a cylindrical hole inclined by 35° to a flat plate at average blowing ratio of M = 0.5 were numerically simulated. The LES results were compared to the experimental data from Seo, Lee and Ligrani (1998), Jung, Lee and Ligrani (2001) and Reynolds-Averaged Navier-Stokes (RANS) results. The experimental results showed that if the oscillation frequency in the main flow was increased, the adiabatic film cooling effectiveness was decreased. The credibility of the LES results relative to the experimental data was demonstrated by the comparison of time-averaged adiabatic film cooling effectiveness, time and phase-averaged temperature contours, contours of Q-criterion, time-averaged velocity profiles, and time and phase-averaged U rms profiles with the RANS results. The adiabatic film cooling effectiveness by LES model showed a good match to the experimental data, while RANS results highly over-predict the centerline effectiveness. Also, the LES results showed more consistent with the experimental data for the time-averaged and phase-averaged temperature contours, time-averaged velocity profiles and time and phase-averaged U rms profiles than the RANS results. RANS did not predict the peak generated by the jet penetration exactly and U rms profiles obtained by RANS approach was much smaller compared to the experimental results. Paper will discuss these results in detail.