The tip-timing technology has been widely developed and has become an industry standard in aircraft engine and gas turbine over past decade. The main application of the tip-timing method is to verify safe operation of blades and monitor the health of blades. But tip-timing technology gets rarely used to the last stage blade of steam turbine. Particularly the blade is designed with an integral shroud, snubber and fir-tree root. The article mainly describes the process of identifying the dynamic characterization of last stage blade with an integral shroud and snubber by contactless measurements provided by tip-timing technology. Attention is focused on the comparison of tip-timing results with the results from strain gauge data. Firstly, the frequency response of the bladed blisk is calculated by using Computer-Aided-Engineering (CAE) technologies. Secondly, according to the results of finite element modal calculation, the location of strain gauge is confirmed. The dynamic strain of blade is measured by utilizing telemetry technology. Finally, according to the design features of integral shroud, the tip-timing probe locations must be accurately confirmed in order to acquire the valid data. All probes are positioned along the radial direction of blades. The rotating vibration test of the bladed blisk has been carried out in the high-speed test rig. In order to validate the tip-timing measurement, all the results from the tip-timing, especially the resonant frequencies and damping ratios, are compared with results from the strain gauges with which only a few blades were equipped.

Single crystal blades used in high pressure turbine bladed disks of modern gas-turbine engines exhibit material anisotropy. In this paper the sensitivity analysis is performed to quantify the effects of blade material anisotropy orientation on deformation of a mistuned bladed disk under static centrifugal load. For a realistic, high fidelity model of a bladed disk both: (i) linear, and (ii) non-linear friction contact conditions at blade roots and shrouds are considered. The following two kinds of analysis are performed: (i) local sensitivity analysis, based on first order derivatives of system response w.r.t design parameters, and (ii) statistical analysis using polynomial chaos expansion. The polynomial chaos expansion is used to transfer the uncertainty in random input parameters to uncertainty in static deformation of the bladed disk. An effective strategy, using gradient information, is proposed to address the “curse of dimensionality” problem associated with statistical analysis of realistic bladed disk.

This paper presents a process for automated start value generation for gas turbine performance simulations using Kriging metamodels. The metamodels are trained on a small number of pre-selected operating points in the flight envelope. Predictions of the trained metamodels are used as initialization parameters for subsequent performance simulations of arbitrary operating points in order to increase robustness and computational speed of the numerical process. Different approaches for the selection of the training points are evaluated. A comparison to the classical approach of table-based initialization is carried out to highlight the advantages and disadvantages of the new methodology. Furthermore, the inclusion of supplementary operating points into the training sample of the metamodels is analyzed. Depending on certain criteria, such as the difference of prediction and simulation result, operating points are included into the sample and a retraining of the metamodels is performed. Simulations using the retrained models as guess value generators are compared to the previous Kriging approach. The advantages and disadvantages of the retraining approach are discussed.

Incorporating recuperators into turbine engines which enables the heat transfer between exhaust gas and compressed air, indicates considerable potential for lower emissions and SFC (specific fuel consumption). However as a matter of fact, they have not yet found wide acceptance in aircraft applications. One of the main potential disadvantages is because in such situations system overall weight needs to be strictly controlled, and there are concerns that the beneficial fuel saving may not offset the additional bulk and weight of the recuperator. In this work, aimed at evaluating the performance of the recuperated rotorcraft and investigating the influence of a recupeator on the whole system, a comprehensive simulation framework has been developed which mainly contains three modules for various flight conditions: Helicopter performance module (HPM), GasTurb engine simulation module and the recuperator weight estimation. Specifically, HPM calculates the helicopter power requirement for different flight conditions (altitude, speed, etc.), and GasTurb is used to simulate the performance of conventional and recuperated turbine engine, and compute the corresponding engine operating point to meet the power demand. The recuperator weight estimation is mainly based on previous studies as a function of mass flow and effectiveness. The methodology is adopted for a typical twin engine light helicopter configuration. In comparison with the conventional non-recuperated cycle, the authors studied the fuel saving potential of the recuperated cycle for different recuperator effectiveness under various flight conditions, and a trade-off analysis was also conducted to identify the flight time required to compensate the additional recuperator weight. The obtained results suggested that the recuperated cycle possesses great potential, especially for long duration and large range mission, but it may not be necessarily suitable for all types of helicopter missions.

Central targets for jet engine research activities comprise the evaluation of improved engine components and the assessment of novel engine concepts for enhanced overall engine performance in order to reduce the fuel consumption and emissions of future aircraft. Since CO 2 emissions are directly related to engine fuel burn, a reduction in fuel consumption leads to lower CO 2 emissions. Therefore improvements in engine technologies are still significant and a multi-disciplinary pre-design approach is essential in order to address all requirements and constraints associated with different engine concepts. Furthermore, an increase in effectiveness of the preliminary design process helps reduce the immense costs of the overall engine development. Within the DLR project PEGASUS (Preliminary Gas Turbine Assessment and Sizing) a multi-disciplinary pre-design and assessment competence of the DLR regarding aero engines and gas turbines was established. The application of modern preliminary design methods allows for the construction and evaluation of innovative next generation engine concepts. The purpose of this paper is to present the developed multi-disciplinary pre-design process and its application to three aero engine models. First, a state of the art twin spool mixed flow turbofan engine model is created for validation purposes. The second and third engine models investigated comprise future engine concepts: a Counter Rotating Open Rotor and an Ultra High Bypass Turbofan. The turbofan used for validation is based on publicly available reference data from manufacturing and emission certification. At first the identified interfaces and constraints of the entire pre-design process are presented. An important factor of complexity in this highly iterative procedure is the intricate data flow, as well as the extensive amount of data transferred between all involved disciplines and among the different fidelity levels applied within the smoothly connected design phases. To cope with the inherent complexity data modeling techniques have been applied to explicitly determine the required data structures of those complex systems. The resulting data model characterizing the components of a gas turbine and their relationships in the design process is presented in detail. Based on the established data model the entire engine pre-design process is presented. Starting with the definition of a flight mission scenario and the resulting top level engine requirements thermodynamic engine performance models are developed. By means of these thermodynamic models, a detailed engine component pre-design is conducted. The aerodynamic and structural design of the engine components are executed using a stepwise increase in level of detail and are continuously evaluated in the context of the overall engine system.

Bearing chamber of a gas turbine engine is generally sealed by pressurized air, separating lubricant from the other zones of the engine. Heat transfer from the wall to air/oil mixture is a challenging engineering problem; predicting heat transfer rate from bearing chamber to oil is important to avoid oil coking and oil fires under high rotational speeds, pressure levels and turbine inlet temperatures. In this study, the inner wall temperature of bearing chamber which is located at the center of front engine structure was investigated numerically. The numerical study involved mainly two thermal modelling methods having two different empirical correlations was performed with finite element solver in order to calculate heat transfer on the wall. First method was based on rotational Reynolds number and Prantl number, in addition to these numbers second one, which is suggested in the literature, is based on oil related and sealing air related Reynolds number, mixture temperature and mixture mass flow. Second approach considers existence of a mixing of gaseous and liquid flow in the core flow unlike first modelling approach. The thermal model was solved by finite element solver and numerical model, assumptions were described with thermal boundary conditions. On the other hand, wall and air thermocouple readings were taken through engine test from the bearing chamber for real engine operating conditions having mainly idle, cruise and maximum power. DN number ranges from 712564 to 2742404, sealing air flow ranges from 46 to 78 g/s and oil flow ranges 22 to 40 g/s for these conditions. The calculated heat transfer coefficients were presented and discussed. The wall temperature predictions of the thermal models, and test measurements were compared. The comparison revealed that analysis results obtained with both correlations were in reasonable agreement with the test. In overall, the second approach predicted metal temperature slightly better at the front support and inner manifold wall, while first approach predicted much better at the rear support wall.

The most progressive liner cooling technology for modern combustion chambers is represented by effusion cooling (or full-coverage film cooling), which is based on the use of several inclined small diameter cylindrical holes. However, as to simulation of the gas turbine combustion chamber, meshing of these discrete holes needs too much computer resource and demanding calculation time. The homogeneous boundary condition was attempted to apply in the throughflow method for the simulation of the full-scale combustion chamber. The verification of this uniform condition was performed through the model of two straight channels. Obtained results were compared with detailed LES simulations, highlighting well accordance and accurate flow structure around the plate. Furthermore, the modelling was used in the simulation of a loop combustion chamber with throughflow method on isothermal state. Performance characteristic and flow fields from this method were then contrasted with the details from the FLUENT simulation upon high geometric fidelity, and prove that the homogeneous boundary condition exerts a good prediction of the performance characteristics and flow field in the combustion chamber.

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

Efforts to achieve NASA’s N+2 and N+3 fuel burn goals have led to various future aircraft concepts. A commonality in all these concepts is the presence of a high degree of interaction among the various disciplines involved. A tightly integrated propulsion/airframe results in distortion in the flow field around the engine annulus. Although beneficial in terms of propulsive efficiency (due to boundary layer ingestion), the impact of distortion on fan performance and operability remains in question for these concepts. As such, rapid evaluation of the impacts of distortion during the conceptual design phase is necessary to assess various concepts. This is especially important given the expansion of the design space afforded by turbo-electric and hybrid-electric distributed propulsion concepts, in which the gas turbine generator and propulsive devices can be decoupled in space. A simple and rapid methodology to assess operability of compressors is the theory of Parallel Compressors (PC). PC theory views the compressor as two compressors in parallel, one with a uniform high P t and the other with a uniform low P t , both operating at the same speed and exiting to a common static pressure. The assumption of two compressors exiting at the common static pressure is not entirely true, especially when the distortion is high. In this paper, the development of a modified parallel compressor model with parametric boundary condition that can capture the impact of non-uniform inflow on fan performance is introduced and validated. Unlike classical PC model, the modified approach introduces a boundary condition dependent on the intensity of distortion (DPCP) at the Aerodynamic Interface Plane (AIP). Additionally, the concept of PC is also extended to Multi-Per Revolution (MPR) distortion. A modeling environment which follows this methodology is created in PROOSIS, an object oriented 0-D cycle code. The model was created using the “compressor” components acting in parallel and a procedure for implementing both design mode and off-design mode solutions was created using the PROOSIS toolset. The example problem was implemented to demonstrate two capabilities — i) the ability of quantifying impacts on thrust and performance of a ducted fan propulsion system, and ii) the ability of predicting loss in stability pressure ratio. The results clearly show the ability of the tool to quantify distortion related losses. The work described in this paper can be integrated to a Multi-Disciplinary Design and Optimization (MDAO) framework along with other disciplines and can be used to evaluate the viability of design space offered by novel aircraft configurations.

Gas turbines in general and aircraft engines in particular undergo frequently dynamic operations. These operations include the routine start-up, load change and shut downs to cover their operation envelope. The frequency of the dynamic operation depends on the size of the engines and the field of application. Engines for commuter aircrafts and particularly helicopter engines operate more often in an off-design mode compared to large commercial aircraft engines and power generation gas turbines. During these routine operations, the compressor mass flow, the pressure ratio, the combustion chamber fuel and air mass flow as well as turbine mass flow change. These changes affect the engine aerodynamic performance and its efficiency. To avoid the inception of rotating stall and surge, high performance gas turbines are equipped with mechanisms that adjust the stator stagger angles thus aligning the stator exit flow angle to the rotor inlet angle, which reduces an excessive incidence. The reduction of incidence angle not only preserves the stable operation of the compressor but it also prevents the compressor efficiency from deterioration. The existence of an inherent positive pressure gradient may cause the boundary layer separation on compressor blades leading to the rotating stall and surge. Such condition, however, does not exist in a turbine, and therefore, there has been no compelling reason to apply the blade adjusting method to the turbine component. For the first time, the impact of turbine blade stagger angle adjustment on the gas turbine efficiency during the operation is shown in this paper. Given a statistically distributed load condition, the extensive dynamic simulation reported in this paper shows how the efficiency can be positively affected through proper blade adjustment. For the time dependent operation, the code GETRAN developed by the author was enhanced to include the turbine blade adjustment as a function of time. To conduct the dynamic simulation with turbine stator stagger angle adjustment during a dynamic operation, the full geometry of the Brown Boveri GT-9 gas turbine was utilized. Starting from the reference stagger angle, it is varied within an incidence range of ± 3 degree. Detailed simulation results show the substantial efficiency improvement through stator stagger blade adjustment.

The development of superior combat aircraft demands the complex integration of the airframe, engine, control system, avionics, and on-board weapon systems. The integration of the engine and the inlet is tantamount to prevailing in an engagement due to the thrust required to execute combat maneuvers. For this reason, test and evaluation methods have been developed to help ensure inlet-engine compatibility by design. The most commonly used methodology characterizes inlet distortion in terms of total-pressure descriptors and correlations. The method includes ground tests employing both wind tunnel and engine test facilities, to acquire the information needed to establish inlet-engine compatibility prior to flight test. Advanced aircraft employing evolving technologies never seen in legacy systems have introduced new challenges to the methodology, and to the ground test methods employed by the methodology. One such challenge arises from the significant flow angularity, or swirl, often found in advanced inlet systems. This paper focuses on the simulation of aircraft inlet swirl during direct-connect turbine engine ground tests. To meet the engine test challenges introduced by advanced aircraft, the Arnold Engineering Development Complex (AEDC) embarked on the development of a swirl generator capable of simulating the different types of swirl expected in future inlet systems over a wide range of swirl angles, and with the ability to remotely set steady-state or transient swirl patterns. The development progressed through a five-step process that culminated in the validation and demonstration of a fully-functional prototype. This paper focuses on the prototype swirl generator and the progression from the establishment of simulation requirements through the prototype validation. Following summaries of each development step, the results of the validation test are presented. The paper also summarizes a recent application of the prototype which not only demonstrated the device in an engine test, but which provided a data set to support swirl methodology development.

Synchro-Self-Shifting (SSS) Overrunning Clutches are used in a myriad of propulsion system configurations for naval and commercial vessels powered by gas turbines and/or combined gas turbine and cruise engines worldwide. Of these, much has been written about high power gas turbine propulsion clutches for large naval vessels (frigates, destroyers, cruisers, etc.), whereas less has been published about the application and experience of the propulsion machinery with Synchro-Self-Shifting Clutches for hydrofoils, hovercraft, fast patrol boats, fast ferries, yachts, etc. Space, weight, and high-speed constraints can be different for high speed gas turbine propulsion systems used in these smaller types of vessels, and can therefore provide gearing challenges, including system design challenges for these clutches. A comparison between Synchro-Self-Shifting overrunning clutches and other types of freewheels will be given discussing the advantages and disadvantages of each, particularly as they relate to high speed gas turbine marine propulsion applications. Lastly, this paper will give some history of a number of high speed gas turbine driven marine propulsion applications with clutches from the early 1960’s until the present, describe various gearing arrangements that were used in particular vessels, articulate where these clutches are incorporated, and discuss the application experience of these clutch installations.

Considering the increasing development of naval destroyers and the obvious advantages of marine gas turbine, this paper designs a novel COGAG (Combined Gas-turbine And Gas-turbine) propulsion system which mainly consists of four GT25 marine gas turbines and one CCG (Cross Connection Gears) for the large destroyer. Firstly, the overall configuration and key devices of COGAG propulsion system are introduced briefly. Then, the typical operating patterns of COGAG propulsion system under different condition are discussed in detail. Finally, many experimental information are further presented. All the results show that the developed COGAG propulsion system not only has higher flexibility and reliability, but also can effectively improve the fuel economy of marine ships.

Compressors are crucial to the efficient operation of a gas turbine; however, fouling, caused by adherence of particles to surfaces, can have a negative effect on compressor performance. In this study, a NASA Stage 35 single-stage axial-flow compressor was employed as the model for numerical simulation using the ANSYS CFX software of the effects caused by mild and severe fouling under salt fog scale. To measure these effects, two distinct models were used. For mild fouling, the simulated stator blade surface roughness was altered nonuniformly; for severe fouling, the simulated stator blade thick-ness was altered. Results indicated that surface roughness caused by mild fouling only has a small effect on compressor performance and no effect on the stable working range. However, changes in the blade thickness as a result of severe fouling have a large effect on compressor performance and a clear effect on the stability of the compressor’s working range. The fouling causes an increase in the boundary layer at the trailing edge of the suction side of the blade thereby increasing the loss of flow; fouling effect after emergence angle, wide design value, and increasing blade of circumferential stress.

The combined diesel-electric and gas turbine (CODLAG) plant is a new type of ship power plant combining the advantages of electric propulsion system and mechanical propulsion system. The requires about ship power grid is lower than full electric propulsion mode, at the same time it can gain the quiet of electric propulsion in the low working conditions and increase the mobility of the ship in the high working conditions. Unlike traditional mechanical propulsion methods and forward-looking all-electric propulsion methods, the CODLAG plant has a coupling between mechanical propulsion and electric propulsion mode. The cooperative working characteristics of two different nature systems is still need further research. For the in-depth research of CODLAG device’s characters, this study built a simulation model of CODLAG device based on Matlab/Simulink and C/C++ platform. A kind of torque-shaft speed double closed-loop control strategy based on PID was used on the CODLAG device. And the typical work condition of CODLAG, including merging, merging off and variable work condition, was simulated in this study. Through the simulation, the dynamic response of main parameters at typical condition have been got. Then, characters of CODLAG device with CPP was simulated at variable condition. And the thrust response was compared with FPP’s. Through comparative analysis, the effectiveness of integrated simulation method specifying to CODLAG device was verified, and some useful conference was provided for the future research.

This paper will discuss the development and testing of an electric drive option designed for the propulsion system of the US Navy’s DDG 51 Class ships. It will briefly explain the history of the Hybrid Electric Drive (HED) program, including that of its predecessor, Proof of Concept (PoC), and the HED’s planned shipboard installation schedule. Operating at lower ship speeds, in a range where the currently installed propulsion gas turbines are less fuel efficient, the HED is expected to increase the ship’s fuel economy, allowing the ship to remain on station accomplishing its mission for a longer period of time. This paper will discuss how the gas turbine propulsion system, in concert with the HED, will be used to provide the most fuel efficient drive combination for various operating scenarios. Also covered will be a description of the major stakeholders involved in the HED’s development and implementation along with some of the constraints and challenges that were encountered in the testing phase of the program, both at the OEM facilities and at the US Navy’s Land Based Engineering Site (LBES) in Philadelphia PA. Planned fuel economy testing results obtained at the LBES facility will also be presented, intended to determine an estimate of the fuel savings that can be expected when the system is first placed in service on USS TRUXTUN (DDG 103) July 2018.

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

The low-pressure turbine has become more important in the last few decades because of the increased emphasis on higher overall pressure and bypass ratios. The desire is to increase blade loading to reduce blade counts and stages in the low-pressure turbine of a gas turbine engine. Increased turbine inlet temperatures for newer cycles results in higher temperatures in the low-pressure turbine, especially the latter stages, where cooling technologies are not used. These higher temperatures lead to higher work from the turbine and this, combined with the high loadings, can lead to flow separation. Separation is more likely in engines operating at high altitudes and reduced throttle setting. At the high Reynolds numbers found at takeoff, the flow over a low-pressure turbine blade tends to stay attached. At lower blade Reynolds numbers (25,000 to 200,000), found during cruise at high altitudes, the flow on the suction surface of the low-pressure turbine blades is inclined to separate. This paper is a study on the flow characteristics of the L1A turbine blade at three low Reynolds numbers (60,000, 108,000, and 165,000) and 15 turbulence intensities (1.89% to 19.87%) in a steady flow cascade wind tunnel. With this data, it is possible to examine the impact of Reynolds number and turbulence intensity on the location of the initiation of flow separation, the flow separation zone, and the reattachment location. Quantifying the change in separated flow as a result of varying Reynolds numbers and turbulence intensities will help to characterize the low momentum flow environments in which the low-pressure turbine must operate and how this might impact the operation of the engine. Based on the data presented, it is possible to predict the location and size of the separation as a function of both the Reynolds number and upstream freestream turbulence intensity (FSTI). Being able to predict this flow behavior can lead to more effective blade designs using either passive or active flow control to reduce or eliminate flow separation.

It is well known that the last stage of a turbine and the subsequent diffuser should be viewed at and designed as a coupled system rather than as single standalone components. The turbine outlet flow imposes the inlet conditions to the diffuser, whereas the recovered dynamic pressure in the diffuser directly controls the turbine back pressure. With changing operating point, the turbine outflow can vary significantly. This results consequently in large variations of the diffuser performance. A major role in the coupled system of turbine and diffuser can be attributed to the tip leakage flow. While it is desirable to minimize the tip leakage with regard to the turbine, a higher leakage mass flow can often be beneficial for the diffuser performance. As there is currently a trend towards aggressive and hence shorter diffusers which are particularly prone to separation, the question arises where the optimum for this tradeoff problem lies. To investigate the performance in the coupled turbine/diffuser system, a generic last stage with shrouded rotor and axial exhaust diffuser have been designed. The components are representative for heavy duty stationary gas turbine applications. Results are presented for three different operating points representing part-load, design-load and over-load condition. Three different seal gap widths are taken into account to control the leakage flow. The results indicate that an operating point dependent optimum gap width can be found for the coupled system efficiency whereas the maximimum turbine performance is always achieved with a minimum gap width.

Endwalls impose a challenge to cool because of the complex system of secondary flows and separation lines disrupting surface film coolant coverage. The interaction of film cooling flows with secondary flow structures is coupled. The momentum exchange of the film coolant with the mainstream affect the formation the secondary flows, which in turn affect the coolant coverage. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics as affected by interactions with coolant flows is required. This study presents experimental and computational results for cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The cascade is subsonic, linear, and stationary with an axisymmetrically-contoured endwall. Two cooling flows are simulated; upstream combustor liner coolant-in the form of an aero-thermal profile simulated in the approach flow and endwall slot film cooling, which is injected immediately upstream of the passage inlet. The experiment is run with engine representative combustor exit flow turbulence intensity and integral length scales, with high turbine passage exit Reynolds number of 1.61 × 10 6 . Measurements are performed with various slot film cooling mass flow rate to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Varying the slot film cooling MFR results in minimal effects on total pressure loss for the range tested. Vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Coolant distribution fields that were previously presented for the same cascade are discussed in context of the aerodynamic measurements. A coolant vorticity parameter presenting the advective mixing of the coolant due to secondary flow vorticity is introduced. This parameter gives developers a new prospective on aerodynamic-thermal performance associated with cooled turbine endwall. The numerical study is conducted for the same test section geometry and is run under the same conditions. The applicability of using RANS turbulence closure models for simulating this type of flow is discussed. The effects of including the combustor coolant in the approach flow is also briefly discussed in context of the numerical results.