Introducing a New Test Rig for Film Cooling Measurements With Realistic Hole Inflow Conditions

Further improvements in film cooling require an in-depth understanding of the influencing parameters. Therefore, a new test rig has been designed and commissioned for the assessment of novel film cooling holes under realistic conditions. The test rig is designed for generic film cooling studies. External hot gas flow as well as internal coolant passage flow are simulated by two individual flow channels connected to each other by the cooling holes. Based on a similarity analysis, the geometry of the test rig is scaled up by a factor of about 20. It furthermore offers the possibility to conduct experiments at high density ratios and realistic approach flow conditions at both cooling hole exit and inlet. The operational range of the new test rig is presented and compared to real engine conditions. It is shown that the important parameters are met and the transfer-ability of the results is ensured. Special effort is put onto the uniformity of the approaching hot gas flow, which will be demonstrated by temperature and velocity profiles. A first measurement of the heat transfer coefficient without film cooling is used to demonstrate the quality of the measurement principle.

Impingement/Effusion Cooling Wall Heat Transfer: Reduced Number of Impingement Jet Holes Relative to the Effusion Holes

Internal wall heat transfer for impingement/effusion cooling was measured and predicted using conjugate heat transfer (CHT) computational fluid dynamics (CFD). The work was only concerned with the internal wall heat transfer and not with the effusion film cooling and there was no hot gas crossflow. Previous work had predicted impingement/effusion internal wall cooling with equal number of holes. The present work investigated a small number of impingement holes and a larger number of effusion holes. The aim was to see if the effusion holes acted as a suction surface to the impingement surface flow and thus enhanced the wall heat transfer. Hole ratios of 1/4, 1/9 and 1/25 were studied by varying the number of effusion holes for a fixed array of impingement holes and a fixed impingement gap, Z, of 8 mm. The Z/D for the impingement holes was 2.7. The impingement hole pitch, X, to diameter, D ratio X/D was 10.6 at a constant effusion hole X/D of 4.7 for all the configurations. The impingement holes were aligned on the midpoint of four effusion holes. The results were computed for a mass flux G from 0.1–0.94 kg/sm2bar for all n. This gave 26 separate CFD/CHT computations. Locally surface, X2, average heat transfer coefficient (HTC), hx, values were determined using the lumped capacitance method. Nimonic 75 metal walls with imbedded thermocouples were used to determine hx from the time constant in a transient cooling experiment following electrical heating to about 80°C. The CHT/CFD predictions showed good agreement with measured data and the highest number of effusion holes for the 1/25 hole ratio gave the highest h. However, comparison with the predicted and experimental results for equal number of impingement and effusion holes for the same Z, showed that there was little advantage of decreasing the number of impingement holes, apart from that of decreasing the Z/D significantly for the 1/15 hole ratio, which increased the heat transfer. The largest number of effusion holes had the highest heat transfer due to the greater internal surface area of the holes and their closer spacing. This was present irrespective of the number of impingement holes and there was no evidence of any benefit of the 25 effusion holes enhancing the single impingement jet heat transfer. For the lowest number of effusion hole there was predicted to be a small disadvantage of reducing the number of impingement jets.

Freestream Flow Effects on Film Effectiveness and Heat Transfer Coefficient Augmentation for Compound Angle Shaped Holes

The use of compound-angled shaped film cooling holes in gas turbines provides a method for cooling regions of extreme curvature on turbine blades or vanes. These configurations have received surprisingly little attention in the film cooling literature. In this study, a row of laid-back fanshaped holes based on an open-literature design, were oriented at a 45-degree compound angle to the approaching freestream flow. In this study, the influence of the approach flow boundary layer thickness and character were experimentally investigated. A trip wire and turbulence generator were used to vary the boundary layer thickness and freestream conditions from a thin laminar boundary layer flow to a fully turbulent boundary layer and freestream at the hole breakout location. Steady-state adiabatic effectiveness and heat transfer coefficient augmentation were measured using high-resolution IR thermography, which allowed the use of an elevated density ratio of DR = 1.20. The results show adiabatic effectiveness was generally lower than for axially-oriented holes of the same geometry, and that boundary layer thickness was an important parameter in predicting effectiveness of the holes. Heat transfer coefficient augmentation was highly dependent on the freestream turbulence levels as well as boundary layer thickness, and significant spatial variations were observed.

Influence of Turbine Blade Leading Edge Profile on Film Cooling With Shaped Holes

This paper presents the turbine blade leading edge model film cooling effectiveness with shaped holes, using the pressure sensitive paint (PSP) mass transfer analogy method. The effects of leading edge profile, coolant to mainstream density ratio and blowing ratio are studied. Computational simulations are performed using the realizable k-ε turbulence model. Effectiveness obtained by CFD simulations are compared with experiments. Three leading edge profiles, including one semi-cylinder and two semi-elliptical cylinders with an after body, are investigated. The ratios of major to minor axis of two semi-elliptical cylinders are 1.5 and 2.0, respectively. The leading edge has three rows of shaped holes. For the semi-cylinder model, shaped holes are located at 0 degrees (stagnation line) and ± 30 degrees. Row spacing between cooling holes and the distance between impingement plate and stagnation line are the same for three leading edge models. The coolant to mainstream density ratio varies from 1.0 to 1.5 and 2.0, and the blowing ratio varies from 0.5 to 1.0 and 1.5. Mainstream Reynolds number is about 100,900 based on the diameter of the leading edge cylinder, and the mainstream turbulence intensity is about 7%. The results provide an understanding of the effects of leading edge profile and on turbine blade leading edge region film cooling with shaped-hole designs.

Turbine Vane Endwall Film Cooling From Cross-Row Configuration With Simulated Upstream Leakage Flow

The performance of a full coverage film cooling configuration called cross-row (CR) configuration including upstream inlet leakage flow was studied by measuring the adiabatic film cooling effectiveness distribution using PSP technique. Experiments were conducted in a blow-down wind tunnel cascade facility at the isentropic exit Mach number of 0.5 corresponding to inlet Reynolds number of 3.8 × 105, based on axial chord length. A free-stream turbulence level was generated as high as 19% with a length scale of 1.7 cm at the inlet. The results are presented as two-dimensional adiabatic film cooling effectiveness distributions on the endwall surface with corresponding spanwise averaged distributions. The focus of this study is to investigate the effect of coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR) on the proposed endwall cooling design. Initially, increased MFR for the endwall cooling and upstream leakage levels up the local adiabatic cooling effectiveness and yields relatively uniform coverage on the entire endwall. However, in either case, highest MFR does not provide any improvement as endwall cooling suffered from the jet lift-off and leakage coolant coverage restricted by the downstream near-wall flow field. Results also indicated a density ratio of 1.5 provides the best performance. Finally, a fair comparison is made with another design called axial-row (AR) configuration from a companion paper.

Experimental Investigation of Numerically Optimized Wavy Microchannels Created Through Additive Manufacturing

The increased design space offered by additive manufacturing can inspire unique ideas and different modeling approaches. One tool for generating complex yet effective designs is found in numerical optimization schemes, but until relatively recently, the capability to physically produce such a design had been limited by manufacturing constraints. In this study, a commercial adjoint optimization solver was used in conjunction with a conventional flow solver to optimize the design of wavy microchannels, the end use of which can be found in gas turbine airfoil skin cooling schemes. Three objective functions were chosen for two baseline wavy channel designs: minimize the pressure drop between channel inlet and outlet, maximize the heat transfer on the channel walls and maximize the ratio between heat transfer and pressure drop. The optimizer was successful in achieving each objective and generated significant geometric variations from the baseline study. The optimized channels were additively manufactured using Direct Metal Laser Sintering and printed reasonably true to the design intent. Experimental results showed that the high surface roughness in the channels prevented the objective to minimize pressure loss from being fulfilled. However, where heat transfer was to be maximized, the optimized channels showed a corresponding increase in Nusselt number.

The Effect of Area Ratio Change via Increased Hole Length for Shaped Film Cooling Holes With Constant Expansion Angles

Shaped film cooling holes are used as a cooling technology in gas turbines to reduce metal temperatures and improve durability, and they generally consist of a small metering section connected to a diffuser that expands in one or more directions. The area ratio of these holes is defined as the area at the exit of the diffuser, divided by the area at the metering section. A larger area ratio increases the diffusion of the coolant momentum, leading to lower average momentum of the coolant jet at the exit of the hole and generally better cooling performance. Cooling holes with larger area ratios are also more tolerant of high blowing ratio conditions, and the increased coolant diffusion typically better prevents jet liftoff from occurring. Higher area ratios have traditionally been accomplished by increasing the expansion angle of the diffuser while keeping the overall length of the hole constant. The present study maintains the diffuser expansion angles and instead increases the length of the diffuser, which results in longer holes. Various area ratios have been examined for two shaped holes: one with forward and lateral expansion angles of 7° (7-7-7 hole) and one with forward and lateral expansion angles of 12° (12-12-12 hole). Each hole shape was tested at numerous blowing ratios to capture trends across various flow rates. Adiabatic effectiveness measurements indicate that for the baseline 7-7-7 hole, a larger area ratio provides higher effectiveness, especially at higher blowing ratios. Measurements also indicate that for the 12-12-12 hole, a larger area ratio performs better at high blowing ratios but the hole experiences ingestion at low blowing ratios. Steady RANS simulations did not accurately predict the levels of adiabatic effectiveness, but did predict the trend of improving effectiveness with increasing area ratio for both hole shapes. Flowfield measurements with PIV were also performed at one downstream plane for a low and high area ratio case, and the results indicate an expected decrease in jet velocity due to a larger diffuser.

Impingement/Effusion Cooling With Low Coolant Mass Flow

A low coolant mass flow impingement/effusion design for a low NOx combustor wall cooling application was predicted, using conjugate heat transfer (CHT) computational fluid dynamics (CFD). The effusion geometry had 4306/m2 effusion holes in a square array with a hole diameter of D and pitch of X and X/D of 1.9. It had previously been shown experimentally and using CHT/CFD to have the highest adiabatic and overall cooling effectiveness for this number of effusion holes. The effect of adding an X/D of 4.7 impingement jet wall with a 6.6 mm impingement gap, Z, and Z/D of 2.0, on the overall cooling effectiveness was predicted for several coolant mass flow rates, G kg/sm2bar. At low G the internal wall heat transfer dominated the overall cooling effectiveness. The addition of impingement cooling to effusion cooling gave only a small increase in the overall cooling effectiveness at all G at 127mm downstream of the start of effusion cooling. An overall cooling effectiveness >0.7 was predicted for a low G of 0.30 kg/sm2bar. This represents about 15% of the combustion air for a typical industrial gas turbine combustor and design changes to reduce this further were suggested based on the predictions of this geometry. The main benefit of the impingement cooling was at the start of the effusion cooling, where the overall cooling effectiveness was dominated by the internal wall impingement and effusion cooling. The separate effusion and impingement cooling were also predicted for comparison with their combination. This showed that the combination of impingement and effusion was not the sum of the individual effusion and impingement heat transfer. The predictions showed that the aerodynamic interactions decreased the effusion and impingement internal wall heat transfer.

Effect of Internal Crossflow Velocity on Film Cooling Effectiveness: Part II — Compound Angle Shaped Holes

In gas turbine engines, film cooling holes are commonly fed with an internal crossflow, the magnitude of which has been shown to have a notable effect on film cooling effectiveness. In Part I of this study, as well as in a few previous studies, the magnitude of internal crossflow velocity was shown to have a substantial effect on film cooling effectiveness of axial shaped holes. There is, however, almost no data available in the literature that shows how internal crossflow affects compound angle shaped film cooling holes. In Part II, film cooling effectiveness, heat transfer coefficient augmentation, and discharge coefficients were measured for a single row of compound angle shaped film cooling holes fed by internal crossflow flowing both in-line and counter to the span-wise direction of coolant injection. The crossflow-to-mainstream velocity ratio was varied from 0.2–0.6 and the injection velocity ratio was varied from 0.2–1.7. It was found that increasing the magnitude of the crossflow velocity generally caused degradation of the film cooling effectiveness, especially for in-line crossflow. An analysis of jet characteristic parameters demonstrated the importance of crossflow effects relative to the effect of varying the film cooling injection rate. Heat transfer coefficient augmentation was found to be primarily dependent on injection rate, although for in-line crossflow, increasing crossflow velocity significantly increased augmentation for certain conditions.

Computational Analysis of a Novel Cooling Scheme for Ultra Compact Combustor Turbine Vanes

The Ultra Compact Combustor (UCC) has the potential to improve overall engine efficiency at a reduced size, while offering a unique component configuration that can allow for novel cooling of turbine vanes located below the circumferential cavity. In this configuration the leading edge of the turbine vane is located upstream of the combustor, with hot gases impinging near 1/5-chord and aft to the trailing edge. Therefore the leading edge remains in a cold region of the flow allowing openings in the stagnation region of the vane to ingest cool core flow. A computational fluid dynamics (CFD) analysis is presented that analyzed stagnation fed cooling variations for the current UCC turbine vane. The variants included several internal only cooling schemes and one external film cooling scheme. The internal only variants showed that with proper component geometry the internal pressure could exceed the external pressure downstream from the leading edge. The film cooling hole variant showed that coolant did in fact exit the vane through 3 of the 8 cooling holes in a similar fashion as a compressor fed scheme. The determination if coolant would exit the holes was highly dependent on the external pressure gradient on the vane.