Area-averaged Heat Transfer Coefficient Research Articles

Effects of the circumferential casing groove configuration on aerothermal performance of a transonic turbine stage

In order to control tip leakage flow and improve the tip aerothermal performance, a circumferential casing groove (CG) configuration is proposed. Taking the typical transonic turbine stage TTM (Thermal Turbomachinery and Machine Dynamics) as the research object, the effects of CG with different heights (b), initial locations (IL) and widths (w) on the aerodynamic loss of the turbine stage and heat transfer characteristics of the tip region are investigated numerically. Results indicate that due to the introduction of CG, the vortex system topology in the tip region changes remarkably. Tip leakage vortex (TLV) is separated into two parts, with TLV-1 being entrained and dissipated by upper passage vortex (UPV) and TLV-2 rebuilding behind the groove and extending to the trailing edge. The interaction between TLV and UPV changes with the development of both. Also, complicated flow-thermal interplays lead to the change of thermal load in the tip region, which attributes to the difference of local turbulence mixing intensity. More quantitatively, a CG with the groove of b = 3.0 g, IL = 0.20 Cax and w = 0.05 Cax can increase ηTS by 0.38%, and reduce the area-averaged heat transfer coefficient (HTC) by 5.28% and 1.49% for the tip and casing surface, respectively, with a slight increase of area-averaged HTC on the blade by 0.29%. Overall, CG with appropriate geometrical parameters can improve aerothermal performance of the turbine tip significantly, without any serious increase of thermal load on the blade.

Non-axisymmetric Endwall film cooling characteristics considering the influences of cylindrical holes and laidback fan-shaped holes

Flow fields near the turbine vane endwall are complicated due to the endwall cross flows. The use of a non-axisymmetric endwall is regarded as an efficient technique to reduce the lateral pressure difference, decreasing the endwall cross flow. Numerical analysis was performed to determine how the non-axisymmetric endwall affected the vortex structure and heat transfer level. The cooling performance was investigated with cylindrical and laidback fan-shaped holes (7–7–7), which were arranged in rows aligned in the axial direction. The results showed that the non-axisymmetric endwall could significantly reduce the circumferential pressure difference and suppress the growth of the passage vortex, and the area-averaged heat transfer coefficient was reduced by 3.34%. The outlet area of the film hole was altered by the non-axisymmetric endwall, and the over-cooled regions may have appeared as a result of the excessive area increase. The influence of the non-axisymmetric endwall was concentrated at 0.4 < Z/Cax < 1.0 for the cylindrical hole. With the increase in M, the film cooling effectiveness of the non-axisymmetric endwall attained a higher level than that of the flat endwall. For the laidback fan-shaped hole, the effect of the non-axisymmetric endwall was confined within 0.25 < Z/Cax < 1.0. The half-period trigonometric function of the non-axisymmetric endwall (HTFN) achieved the optimal cooling performance for three blowing ratios. However, the periodic trigonometric function of the non-axisymmetric endwall (PTFN) only outperformed the flat endwall when M= 1.5.

Effect of wear damage on aero-thermal performance of the film-cooled squealer tip in a turbine stage

Effects of wear damage on the aerodynamic performance, heat transfer and film cooling effect in the squealer tip region of a gas turbine stage were investigated under engine condition. Two wear parameters, including the starting-location of wear ( sl) and wear depth ( wd), were introduced to describe the degree of wear damage. Three starting-locations of wear ( sl = 32% C ax, 58% C ax and 79% C ax) and three wear depths ( wd = 0.33 mm, 0.67 mm and 1.0 mm) were selected to compute the heat transfer coefficient and film cooling effectiveness on the squealer tip at four blowing ratios ( M = 0.25, 0.5, 0.75 and 1.0). The pressure loss and isentropic efficiency of the turbine stage were calculated to assess the aerodynamic performance degradation induced by the tip wear. The results indicate that the pressure loss in the blade passage near tip region is increased and the stage efficiency is decreased with the increase of wd. Compared to the original design case at M = 0.75, the pressure loss in the rotor blade passage is increased by 7.39% and the stage efficiency is decreased by 2.22% at most after wear. As wd increases, the area-averaged heat transfer coefficient on the cavity floor is increased and film cooling effectiveness is decreased. For the worn cases at M = 0.25, 0.5, 0.75 and 1, the area-averaged heat transfer coefficient on the cavity floor is increased respectively by 87.7%, 62.3%, 71.3% and 88.0% at most, and the area-averaged film cooling effectiveness is decreased respectively by 37.2%, 26.8%, 26.7% and 22.2% at most, as compared to the original design cases.

Modifications of double-rim geometry to improve thermal performance of squealer tip in a turbine stage

The double-recessed (i.e. double-rim) design is an alternative to improve thermal performance of the squealer tip. In this paper, the aero-thermal performance of double-rim squealer tip in a turbine stage was numerically investigated under engine condition. The heat transfer and film cooling effect between the double-rim squealer tip and conventional squealer tip were compared to show the superior thermal characteristic of double-rim design. The interactions between the leakage flow and coolants were analyzed to address the heat transfer reduction and film cooling improvement for the double-rim configuration. To further reduce the thermal load on double-rim squealer tip, the height and arc length of the inner rim were modified to increase the coolant accumulations in squealer cavities. The results show that the tip and pressure side cooling flow is well attached on the cavity floor and suction side rim of double-rim squealer tip as compared to the conventional squealer tip. With the proper arrangement of inner rim height and arc length, the modified double-recessed squealer tip helps to reduce the area-averaged heat transfer coefficient by 90.3%, and improve the film cooling effectiveness by 49.6%, while the isentropic stage efficiency is almost not affected, as compared to the conventional squealer tip.

가스터빈의 연소기 냉각을 위한 이중벽 냉각 구조에서 유출 구멍의 복합각도 변경에 따른 냉각 성능 변화에 대한 수치해석적 연구

For the cooling of the combustor liner, this investigation considers a double-wall cooling structure in which the compound angle and orientation angle of the effusion hole are 30° and ±30°. respectively. Also, to analyze the effect of orientation angle, data for comparison in a simple angle configuration with an orientation angle of 0° is provided. In this study, the heat transfer and flow characteristics of the double wall cooling structure by the change in orientation angle were numerically analyzed. The results of the numerical simulations was obtained using ANSYS FLUENT V 21.1 numerical codes with a shear stress transfer k-ω turbulence model. Data are provided for main flow Reynolds numbers from 142,000 to 149,000. and overall blowing ratios from 5.5 to 7.4. The use of compound and orientation angle arrangements inhibits the formation of horseshoe-shaped vortices,BRresulting in asymmetry of the counter rotating vortex. As a result, the lift-off of the coolant injected from the effusion holes is suppressed from the effusion cooling plate. When orientation angle is changed from 0° to ± 30°, the area-averaged film cooling effectiveness and area-averaged heat transfer coefficient increase by 7.7% and 12.4% overall, respectively. The skewed and non-symmetric shape of the horseshoe-shaped vortex is also increased by the use of orientation angle arrangement, which also provide localized improvements in surface thermal protection.

Comparisons of cooling performance and flow characteristics of a combustor liner plate with compound angle and simple angle effusion holes

The present investigation considers a unique compound angle arrangement, never previously considered, with α = 30°, β = ±30°, such that the compound angle of the effusion holes changes from β = +30° to β = −30°, as adjacent rows of effusion holes are encountered. Also provided are data for comparison from a simple angle configuration with α = 30°, β = 0°. Coolant is supplied to both effusion hole arrays using arrays of impingement cooling jets. The results of the numerical simulations are obtained using the ANSYS FLUENT V 21.1 numerical code, with a shear stress transport (SST) k–ω turbulence model. Data are provided for main flow Reynolds numbers from 142,000 to 155,000, impingement flow Reynolds numbers from 7900 to 18,000, effusion flow Reynolds numbers from 10,400 to 23,600, and overall blowing ratios from 3.3 to 7.4. When β is changed from 0° to ±30°, area-averaged film cooling effectiveness values and area-averaged heat transfer coefficient magnitudes increase overall by approximately 2.89% and 11.7%, respectively. This is because the use of compound angle arrangements often inhibits the formation of the horseshoe-shaped vortex, and the resulting counter-rotating vortex legs, that often promote lift off of significant film concentrations from the test surface. The skewed and non-symmetric character of the horseshoe-shaped vortex is also increased by the use of compound angle arrangement, which also sometimes provides local improvements in surface thermal protection.

Thermal performance improvement on winglet tip of a turbine stage under engine condition-part II: Ejection angle adjustment

Part I has indicated that the vertical ejection causes pronounced coolant lifting-off effect on the winglet tip of turbine stage. To solve this problem, Part II presents the study on adjusting ejection angle to reduce thermal load on the winglet tip under engine condition. Three procedures were taken to realize the mitigation of thermal load on the winglet tip of turbine stage. Firstly, the streamwise ejection angle was varied to determine a suitable ejection angle for the majority of cooling holes. Secondly, different numbers of tip cooling holes were selected to adjust the ejection angles towards P.S. (pressure side), aiming at enhancing the film cooling effect at the front part of the cavity floor. Lastly, different P.S. deflection angles were tested to find the suitable streamwise ejection angles for the selected holes. The results show that the inclined coolant ejections are quite effective to reduce the thermal load on the winglet tip. As the P.S. defection angles for the tip cooling holes at the front part of the cavity floor equal to 60°, and the streamwise ejection angles for the other cooling holes are set to 120°, the lowest thermal load is achieved on the winglet tip of turbine stage. After adjustment, the area-averaged heat transfer coefficient on the tip surface is decreased by 38.84%, and the area-averaged film cooling effectiveness is increased by 42.24%, as compared with the vertical ejection case. Moreover, the uniformity of the coolant coverage on the winglet tip has been improved, significantly.

Effect of unsteady passing wake on the aerodynamic and heat transfer performance of the turbine blade squealer tip

By solving the Unsteady Reynolds-averaged Navier-Stokes equations and standard k-ω turbulence model, effect of unsteady passing wake on squealer tip heat transfer performance for a gas turbine stage with different rotating speed is numerically investigated. The results of turbulence model validation are in good agreement with the experimental data. The grid independence verification is also satisfied. The results indicate that the tip area-averaged heat transfer coefficient fluctuates periodically with time, and a “platform stage” is existed within the heat transfer curve from trough to peak. The heat transfer curve is divided into two independent phases by the “platform stage”. In the first phase, the heat transfer variation is concentrated in the tip front part, and the average heat transfer coefficient at this region increases by 3.4%. In the second phase, the variation of heat transfer in the tip rear part is obvious, and the average heat transfer coefficient of this region increases by 5.1%. Compared with the first and second phases, the heat transfer coefficient curve in the third phase shows the characteristics of uniform decline. The existence of the “platform stage” is closely related to the blade rotating speed. The “platform stage” will disappear as the blade rotating speed deviates from the standard condition.

Investigations into heat transfer and film cooling effect on winglet-squealer tips in a turbine stage

Heat transfer and film cooling effect on three kinds of winglet-squealer tips (i.e., conventional squealer tip, no P.S. rim tip and P.S. winglet tip) in a turbine stage are numerically investigated at two operation conditions (i.e., laboratory testing condition and sea-level take-off condition) and three cooling hole arrangements (i.e., un-cooled, with tip holes, and with both tip and pressure side holes), and the numerical results are also compared with those in stationary cascades. To reduce the heat transfer while enhancing film cooling effect on the P.S. winglet tip, ejection angles for the tip and pressure side cooling holes are varied to reveal the superior performance of P.S. winglet tip in turbine stage. The results show that the heat transfer coefficient and film cooling effectiveness distributions on three kinds of tips exhibit different trends in the stationary and rotating cases. Tip rotating significantly affects the trajectories of cooling flow in the conventional squealer tip and no P.S. rim tip due to different vortex system induced in the cavity. However, the relative rotating has little influence on the flow structures in the P.S. winglet tip. In un-cooled condition, the area-averaged heat transfer coefficient on P.S. winglet tip is lower than the conventional squealer tip by 11.6%. For the vertical ejection case in turbine stage, the P.S. winglet tip has a slightly worse heat transfer and film cooling effect than the conventional squealer tip. However, with inclined ejections, the area-averaged heat transfer coefficient on winglet tip in turbine stage is lower than the conventional squealer tip by 15.3%, and the area-averaged film cooling effectiveness on winglet tip is higher than the conventional squealer tip by 12.1%.

Investigations Into Heat Transfer and Aerodynamic Performance of a Worn Squealer Tipped Turbine Stage

Abstract Heat transfer and aerodynamic performance in worn squealer tip gap of a high-pressure gas turbine stage were numerically investigated. Effects of the starting location of wear and wear depth on tip heat transfer coefficient distributions and stage efficiency were analyzed to evaluate the aero-thermal performance degradations in the gas turbine stage after wear. At three starting locations of wear and five wear depths, flow patterns in worn squealer tip gap of the turbine stage were visualized and compared with the original design case. The results show that the counter-rotating vortex systems in tip cavity, as well as the interactions between leakage vortex and passage vortex, are significantly affected by the degree of wear damage. The starting location of wear and wear depth have pronounced influences on heat transfer and aerodynamic performance in squealer tip gap. After wear, the stage efficiency is decreased by about 0.3–1%, as the wear depth is equal to clearance gap size. In the serious worn case, thermal load on tip cavity floor is increased by about 60%, while the heat transfer on rims is reduced by about 20%. However, compared with the original design case, the area-averaged heat transfer coefficient on shroud is reduced by 5% at most.

Effects of the novel rib layouts on the tip leakage flow pattern and heat transfer performance of turbine blade

The steady simulation with flow field and heat transfer performance for different rib layouts is investigated. By referring the GE-E3 turbine blade tip (Case 1) profile as a prototype model, four kinds of rib layouts of full rib structure (Case 2), half rib structure connected with suction side (Case 3), half rib structure connected with pressure side (Case 4), and half rib structure in the rear squealer cavity (Case 5) was designed. The availability of k-ω turbulence model was validated through comparison of heat transfer coefficient distribution with the experimental data. The area-averaged heat transfer coefficients of Case 5 decreases up to 11.3%, 3.1%, 11.3%, and 2.8% in comparison to Cases 1, 2, 3, and 4. The total pressure coefficient of Case 5 reduces by 1.4%, 2.7%, and 4.0% compared to Cases 1, 2, and 3. The results reveal that, among the five novel rib layout cases, Case 5 obtains the optimal comprehensive performance under the performance of considering tip heat transfer and cascade loss. Further investigation on the influence of Cases 1 and 5 at different rotational speeds on the flow field and heat transfer characteristics of the tip clearance is also illustrated and discussed.

Computational appraisal of the thermalhydraulic charac-teristics of supercritical carbon dioxide in heated mini-channel for HVAC applications

Over last couple of decades, supercritical fluid has found enhanced applications in numerous commercial sectors, encompassing aerospace, chemical, fast reactors, fusion rectors and renewable energy. Supercritical fluid is of major interest to the thermal engineers, primarily owing to its superior heat transport characteristics around the pseudocritical temperature. Supercritical carbon dioxide has particularly been identified as the next-generation coolant in power industry, due to its low critical temperature and reasonable critical pressure. The heating, ventilation and air-conditioning sector has also seen increasing popularity of supercritical CO2 with the phasing out of the conventional halocarbons, as it has zero ozone depletion potential and substantially smaller global warming potential. While decent volume of literature is available regarding the application of supercritical CO2 in macrochannel, thermalhydraulics of supercritical CO2 in minichannel lacks comprehensive investigation, and the present work attempts to fill that specific void through a computational appraisal of a heated horizontal minichannel of 2 mm inner diameter. Systematic simulations have been performed to explore the role of the wall heat flux ranging from 30–50 kW/m2, operating pressure ranging from 8–9 MPa, inlet temperature ranging from 295–315 K and buoyancy parameter on the thermalhydraulic characteristics of supercritical CO2. The Reynolds number in the present simulation ranges from 10000-17000. This article is aimed to explore the phenomena of heat transfer deterioration in horizontal minichannel subjected to uniform wall heat flux. It is observed that increase in wall heat flux leads to lower area-averaged heat transfer coefficient The results show an early heat transfer deterioration on top half surface whereas, normal heat transfer is observed on bottom half surface. Heat flux has significant effect on heat transfer deterioration and higher heat flux leads to reduction in peak value of heat transfer coefficient. At higher inlet temperature heat transfer coefficient decreases and peak of heat transfer coefficient shifted in the upstream direction. Also, peak of heat transfer coefficient vanishes for inlet temperature higher than pseudocritical temperature due to non-significant variation in thermophysical properties at higher temperature. Stratification of temperature, velocity and density is observed along the channel due to local buoyancy produced by non-linear thermophysical property variation of supercritical CO2.

Influence of wear damages on aerodynamic and heat transfer performance in squealer tip gap

Numerical simulations were performed to investigate the effect of rim wear on aerodynamic and heat transfer performance in a squealer tip gap of gas turbine. Flow patterns, heat transfer coefficient distributions, and total pressure losses in tip gap regions were obtained at three starting-locations of wear (sl = 25%Cax, 50%Cax and 75%Cax. Cax: axial chord) and five wear-depths (wd = 1 mm, 2 mm, 3 mm, 4 mm and 5 mm). With the existing experimental data, numerical methods were carefully examined with respect to the turbulence model and mesh density. Results indicate that flow fields in squealer tip gap are significantly affected by the geometrical parameters of worn rims. As wd (wear depth) increases, high heat transfer area on cavity floor near leading edge is enlarged, whereas heat transfer coefficient on worn rims are decreased. The intensity of leakage vortex is increased significantly while the intensity of passage vortex is decreased, leading to increased total pressure loss in tip region. As sl (starting-location of wear) moves towards the leading edge, total pressure loss in cascade is increased, and area-averaged heat transfer coefficient on cavity floor is increased sharply. However, the area-averaged heat transfer coefficient on rims is decreased, significantly.

Effects of the squealer winglet structures on the heat transfer characteristics and aerodynamic performance of turbine blade tip

The heat transfer characteristics and aerodynamic performance of three different squealer winglet structures based on GE-E3 turbine rotor squealer tip are numerically investigated, which is compared with the conventional squealer tip. The Reynolds-Averaged Navier-Stokes (RANS) solutions and standard k-ω turbulence model are solved to analyze the aerothermal performance of rotor tip with different squealer winglet structures. The numerical method is verified with experimental data. The numerical results show that the area-averaged heat transfer coefficient of squealer winglet structure at the pressure side, suction side and both the pressure and suction side are respectively reduced by 12.2%, 17.1% and 19.8% compared to the conventional squealer tip. The squealer winglet structure on the blade tip mostly influence the flow structures inside the cavity with the decreased strength of the pressure side corner vortex and the scraping vortex, which reduce the heat transfer coefficient and the leakage flow rate of the blade tip. The squealer winglet structure at the pressure side is found to increase the total pressure loss coefficient of the rotor blade by 8.5% in contrast to the conventional squealer tip. However, the squealer winglet structure at the suction side and both sides effectively reduce the total pressure loss coefficient by 8.5% and 2.5% respectively. The results reveal that squealer winglet structure at the suction side has the tradeoff of the heat transfer characteristics and aerodynamic performance among four different turbine rotor squealer tip profiles.

Quench subcooled-jet impingement boiling: Staggered-array jets enhancement

In quench subcooled-jet boiling, an effective cooling of large surface area is achieved with jet arrays. Here a stainless-steel plate initially superheated at 900 °C is cooled with seven hexagonally arranged water [1 atm, 20 °C (80 °C subcooling)] jets (7j), and with varying jet separation distance (with a jet Reynolds number ReD = 5000). Using high-definition visualization synchronized with the surface thermal characterization based on the inverse-conduction analysis allows calculation of the spatial and temporal variations of the local and average heat flux (q) and heat transfer coefficient (h). The local and temporal peaks in q and h occur under the jets with enhancements in the regions of their interactions. We compare these results with the prior two jets (2j) and single jet (1j) results (with ReD = 15,000). For a jet separation of four nozzle diameter (S/Dn = 4), the 7j behavior is close to the single jet (1j), while with larger separation extra h peaks occur and the cooling becomes more effective (larger cooling area, but with smaller area-averaged heat transfer coefficient 〈h〉). For S/Dn = 10, the 7j enhanced cooling effectiveness is smaller than the 2j effectiveness, suggesting using similar total liquid flow rate, fewer jets are preferred. The cooling effectiveness enhancement for the S/Dn = 6 arrangement is the highest, reaching 50% and covering a large high-h area, and over a long elapsed time.

Numerical analysis of the aerodynamic performance & heat transfer of a transonic turbine with a partial squealer tip

The purpose of this study is to evaluate thermal load and aerodynamic performance by applying different shapes of tips depending on the velocity region in a transonic turbine where subsonic and supersonic regions coexist. The RS1S, which is the transonic turbine of SNECMA, is used. To understand the tendencies, geometries A, B and C were created by varying the length of the pressure surface rim and the suction surface rim, and they were compared with a flat tip, squealer tip, pressure surface rim and suction surface rim. RANS computational analysis was performed using the commercial code ANSYS CFX 18.0, and aerodynamic performance was evaluated by using the mass-averaged total pressure loss coefficient at the 133% Cax and the mass flow rate of tip leakage. Thermal load was also evaluated by using the area-averaged heat transfer coefficient at the tip. Although the mass flow rate of leakage was smallest with the squealer tip, geometry C, the 75% rim of the pressure surface and the suction surface showed the lowest total pressure loss coefficient, which was 6.5% lower than that of the flat tip. In addition, when only the suction surface rim was applied, the area-averaged heat transfer coefficient was 10% lower than that of the flat tip.

Length to diameter ratio effect on heat transfer performance of simple and compound angle holes in thin-wall airfoil cooling

Heat transfer coefficients on a flat plate surface downstream a row of simple and compound angle cylindrical holes are investigated using high-resolution thermographic liquid crystal technique. A variation of flow parameters including blowing ratio, and geometry parameters including compound angle and length-to-diameter ratio are examined. Blowing ratios (M) ranging from 0.3 to 2, length to diameter ratios (L/D) from 0.5 to 5, and two compound angle (β: 0°, 45°) are employed composing a test matrix of 70 test cases. Detailed local, spanwise averaged, and area averaged heat transfer coefficients hf/h0 are presented to illustrate the effect of length-to-diameter ratio and compound angle. The film cooling performance is also evaluated using NHFR method and Δφ method by combining adiabatic film effectiveness and heat transfer coefficient data. Results indicate that Δφ method has superiority in evaluating film cooling performance due to its direct reflection of temperature reduction by film protection.

Thermal Management of a Transonic Turbine: Leakage Flow and Endwall Contouring Effects

Comparison of heat transfer performance of a nonaxisymmetric contoured endwall to a planar baseline endwall in the presence of leakage flow through a stator-rotor rim seal interface is reported in this paper. Heat transfer experiments were performed on a high turning turbine airfoil passage at a transonic exit Mach number at different leakages to mainstream mass flow ratios. The contoured endwall geometry was optimized for secondary flow aerodynamic losses. Transient infrared thermography was used to measure the endwall surface temperature. Simultaneous calculation of the heat transfer coefficient (HTC) and adiabatic cooling effectiveness (ETA) was performed assuming one-dimensional semi-infinite transient conduction. Results show nominal reduction in the area-averaged HTC but significantly higher coolant film coverage using the contoured endwall compared to baseline endwall at different levels of coolant MFR. Interestingly, significant reduction in the area-averaged heat transfer coefficient was observed for cases without coolant flow in the presence of a purge slot when compared with a solid cascade without any cooling features. The backward-facing step seemed to have a first-order impact on endwall heat transfer coefficient distribution compared to coolant blowing cases. The contoured endwall also showed major improvement in net heat flux reduction (combining the effects of the HTC and ETA) especially near the suction side of the platform.

CFD simulation of single-phase heat transfer in a rotor-stator spinning disc reactor

A computational fluid dynamics (CFD) study on hydrodynamics and heat transfer in a rotor-stator spinning disc reactor (RSSDR), an inventive reactor technology engineered within the frame of process intensification, was performed with OpenFOAM®. This allows the investigation of locally resolved temperature profiles and heat transfer coefficients in the reactor gap, which is experimentally difficult for RSSDRs with small dimensions, while keeping computational costs low. The retrieved radial and tangential velocity and temperature profiles are in agreement with literature. Local stator-side heat transfer coefficients were determined, which increase with higher disc speeds and with increasing throughputs. Area-averaged heat transfer coefficients across the reactor gap were calculated and validated against experimental results performed on the setup simulated in this study; the numerical and experimental data are in qualitative agreement. This study presents the possibility to characterize an RSSDR in terms of hydrodynamics and heat transfer properties by means of a two-dimensional axisymmetric geometry and Reynolds-averaged Navier-Stokes (RANS) turbulence modeling in OpenFOAM®. Expensive measurements, which are difficult to perform particularly for RSSDRs with small dimensions, could thus be avoided, simulating industrial applications at the conditions of interest and designing the RSSDR setup accordingly.

Development of a heat sink module for a near-term DEMO divertor

A heat sink module has been developed for high-heat flux applications in the divertor region of a near term DEMO divertor. The concept employs water jet impingement, in a cascading flow path, together with CuCrZr as the structural material. This fluid-flow architecture maximises the heat transfer coefficient obtained for a given coolant flow rate. Combined with high operational pressures, it facilitates increased coolant outlet temperatures. A conjugate CFD study was performed to compare its performance to the ITER divertor under a representative heat flux boundary condition of 10 MW/m2. A transient thermochromic liquid crystal technique was used to experimentally validate the CFD heat transfer data, using air as the working fluid at matched Reynolds numbers. These results showed that area averaged heat transfer coefficients could be accurately predicted by the CFD to within ∼12%.