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NASA NTRS2003

Film Cooling Flow Effects on Post-Combustor Trace Chemistry

Film cooling injection is widely applied in the thermal design of turbomachinery, as it contributes to achieve higher operating temperature conditions of modern gas turbines, and to meet the requirements for reliability and life cycles. It is a significant part of the high-pressure turbine system. The film cooling injection, however, interacts with the main flow and is susceptible to have an influence on the aerodynamic performance of the cooled components, and through that may cause a penalty on the overall efficiency of the gas turbine. The main reasons are the loss of total pressure resulting from mixing the cooling air with mainstream and the reduction of the gas stagnation temperature at the exit of the combustion chamber to a lower value at the exit of nozzle guide vane. In addition, the impact of the injected air on the evolution of the trace species of the hot gas is not yet quite clear. This work computationally investigates the film cooling influence on post-combustor trace chemistry, as trace species in aircraft exhaust affect climate and ozone.

Wey, Thomas↗

NASA NTRS2000

Unsteady High Turbulence Effects on Turbine Blade Film Cooling Heat Transfer Performance Using a Transient Liquid Crystal Technique

Unsteady wake effect, with and without trailing edge ejection, on detailed heat transfer coefficient and film cooling effectiveness distributions is presented for a downstream film-cooled gas turbine blade. Tests were performed on a five-blade linear cascade at an exit Reynolds number of 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.4 to 1.2; air and CO2 were used as coolants to simulate different density ratios. Surface heat transfer and film effectiveness distributions were obtained using a transient liquid crystal technique; coolant temperature profiles were determined with a cold wire technique. Results show that Nusselt numbers for a film cooled blade are much higher compared to a blade without film injection. Unsteady wake slightly enhances Nusselt numbers but significantly reduces film effectiveness versus no wake cases. Nusselt numbers increase only slic,htly but film cooling, effectiveness increases significantly with increasing, blowing ratio. Higher density coolant (CO2) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher 0 effectiveness at lower blowing ratios (M = 0.8). Trailing edge ejection generally has more effect on film effectiveness than on the heat transfer, typically reducing film effectiveness and enhancing heat transfer. Similar data is also presented for a film cooled cylindrical leading edge model.

Han, J. C.↗

NASA NTRS1954

Comparison of effectiveness of convection-, transpiration-, and film-cooling methods with air as coolant

Various parts of aircraft propulsion engines that are in contact with hot gases often require cooling. Transpiration and film cooling, new methods that supposedly utilize cooling air more effectively than conventional convection cooling, have already been proposed. This report presents material necessary for a comparison of the cooling requirements of these three methods. Correlations that are regarded by the authors as the most reliable today are employed in evaluating each of the cooling processes. Calculations for the special case in which the gas velocity is constant along the cooled wall (flat plate) are presented. The calculations reveal that a comparison of the three cooling processes can be made on quite a general basis. The superiority of transpiration cooling is clearly shown for both laminar and turbulent flow. This superiority is reduced when the effects of radiation are included; for gas-turbine blades, however, there is evidence indicating that radiation may be neglected.

Eckert, E R G↗

NASA NTRS1962

EVAPORATIVE FILM COOLING FOR HYPERSONIC REENTRY VEHICLES

Evaporative film cooling on blunt nose models has been investigated in a 12″ × 12″ hypersonic blowdown wind tunnel at Mach number 7 with stagnation temperatures up to 2500 °R. Water as coolant was supplied continuously through orifices located near the stagnation point of a hemisphere-cylinder body (3 inch diameter) and a blunt cone with 10° half angle (base diameter 5.5 inches) forming an evaporating water film covering the body. This results in nearly uniform wall temperatures about 30 °R below the local boiling temperature of the water.

SPACE SCIENCE↗

NASA NTRS1974

An adverse effect of film cooling on the suction surface of a turbine vane

Film-cooling-air ejection from the suction surface of a turbine vane was investigated. This investigation was conducted in a four-vane cascade on a J75 size turbine vane which had a row of holes near the leading edge. The experimental data are presented. It was found that a small amount of film-cooling air has a detrimental effect on the downstream vane wall cooling effectiveness. It was also shown that the presence of the film-cooling holes, without blowing, also cause an increase in vane wall temperatures. These results came from an increase in the gas-side heat transfer coefficient that was apparently caused by a laminar or transitional boundary layer becoming transitional or turbulent.

Gladden, H. J.↗

NASA NTRS2003

Active Control of Jets in Cross-Flow for Film Cooling Applications

Jets in cross-flow have applications in film cooling of gas turbine vanes, blades and combustor liners. Their cooling effectiveness depends on the extent to which the cool jet-fluid adheres to the cooled component surface. Lift-off of the cooling jet flow or other mechanisms promoting mixing, cause loss of cooling effectiveness as they allow the hot "free-stream" fluid to come in contact with the component surface. The premise of this project is that cooling effectiveness can be improved by actively controlling (e.9. forcing, pulsing) the jet flow. Active control can be applied to prevent/delay lift-off and suppress mixing. Furthermore, an actively controlled film-cooling system coupled with appropriate sensory input (e.g. temperature or heat flux) can adapt to spatial and temporal variations of the hot-gas path. Thus, it is conceivable that the efficiency of film-cooling systems can be improved, resulting in coolant fluid economy. It is envisioned that Micro Electro-Mechanical Systems (MEMS) will play a role in the realization of such systems. As a first step, a feasibility study will be conducted to evaluate the concept, identify actuation and sensory elements and develop a control strategy. Part of this study will be the design of a proof-of-concept experiment and collection of necessary data.

Nikitopoulos, Dimitris E.↗

NASA NTRS2008

A Numerical Study of Anti-Vortex Film Cooling Designs at High Blowing Ratio

A concept for mitigating the adverse effects of jet vorticity and liftoff at high blowing ratios for turbine film cooling flows has been developed and studied at NASA Glenn Research Center. This "anti-vortex" film cooling concept proposes the addition of two branched holes from each primary hole in order to produce a vorticity counter to the detrimental kidney vortices from the main jet. These vortices typically entrain hot freestream gas and are associated with jet separation from the turbine blade surface. The anti-vortex design is unique in that it requires only easily machinable round holes, unlike shaped film cooling holes and other advanced concepts. The anti-vortex film cooling hole concept has been modeled computationally for a single row of 30deg angled holes on a flat surface using the 3D Navier-Stokes solver Glenn-HT. A modification of the anti-vortex concept whereby the branched holes exit adjacent to the main hole has been studied computationally for blowing ratios of 1.0 and 2.0 and at density ratios of 1.0 and 2.0. This modified concept was selected because it has shown the most promise in recent experimental studies. The computational results show that the modified design improves the film cooling effectiveness relative to the round hole baseline and previous anti-vortex cases, in confirmation of the experimental studies.

Heidmann, James D.↗

NASA NTRS1982

Film-cooling effectiveness with developing coolant flow through straight and curved tubular passages

The data were obtained with an apparatus designed to determine the influence of tubular coolant passage curvature on film-cooling performance while simulating the developing flow entrance conditions more representative of cooled turbine blade. Data comparisons were made between straight and curved single tubular passages embedded in the wall and discharging at 30 deg angle in line with the tunnel flow. The results showed an influence of curvature on film-cooling effectiveness that was inversely proportional to the blowing rate. At the lowest blowing rate of 0.18, curvature increased the effectiveness of film cooling by 35 percent; but at a blowing rate of 0.76, the improvement was only 10 percent. In addition, the increase in film-cooling area coverage ranged from 100 percent down to 25 percent over the same blowing rates. A data trend reversal at a blowing rate of 1.5 showed the straight tubular passage's film-cooling effectiveness to be 20 percent greater than that of the curved passage with about 80 percent more area coverage. An analysis of turbulence intensity detain the mixing layer in terms of the position of the mixing interface relative to the wall supported the concept that passage curvature tends to reduce the diffusion of the coolant jet into the main stream at blowing rates below about. Explanations for the film-cooling performance of both test sections were made in terms differences in turbulences structure and in secondary flow patterns within the coolant jets as influenced by flow passage geometry.

Papell, S. S.↗

NASA NTRS1977

Experimental flow coefficients of a full-coverage film-cooled-vane chamber

Ambient- and elevated-temperature flow tests were performed on a four-times-actual-size model of an impingement- and film-cooled segment of a core engine turbine vane. Tests were conducted with the impingement and film cooling plates combined to form a chamber and also with each of the individual separated plates. For the combined tests, the proximity of the film cooling plate affected the flow of coolant through the impingement plate, but not conversely. Impingement flow is presented in terms of a discharge coefficient, and the film cooling flow discharging into still air with no main stream gas flow is presented in terms of a total pressure-loss coefficient. The effects of main stream gas flow on discharge from the film cooling holes are evaluated as a function of coolant to main-stream gas momentum flux ratio. A smoothing technique is developed that identifies and helps reduce flow measurement data scatter.

Meitner, P. L.↗

DOE OSTI2022

Predictive Modeling of Local Film-Cooling Flow on a Turbine Rotor Blade

Abstract In the turbine section of a modern gas turbine engine, components exposed to the main gas path flow rely on cooling air to maintain hardware durability targets. Therefore, monitoring turbine cooling flow is essential to the diagnostic and prognostic efficacy of a condition-based operation and maintenance (CBOM) approach. This study supports CBOM goals by leveraging supervised machine learning to estimate relative changes to local film-cooling flowrate using surface temperature measured on the pressure side of a rotating turbine blade operating at engine-relevant aerothermal conditions. Throughout the lifetime of a film-cooled turbine component, characteristics of the film-cooling flow—such as film trajectory and cooling effectiveness—vary as degradation-driven geometry distortions occur, which ultimately affects the relationship between the model input and the model output—film-cooling flowrate predictions. The present study addresses this complication by testing a data-driven model on multiple turbine blades of the same nominal design, but with each blade exhibiting different localized film-cooling flow characteristics. By testing the model in this manner, strategies for mitigating the detrimental effects of film-cooling flow characteristic variations on model performance were investigated, and the corresponding flowrate prediction accuracy was quantified.

Engineering↗

NASA NTRS1999

A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

A three-dimensional Navier-Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Glenn Research Center experiment and has both circular cross-section and shaped film cooling holes. This complex geometry is modeled using a multi-block grid which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges.

Heidmann, James D.↗

NASA NTRS1976

Heat-Transfer Characteristics of Partially Film Cooled Plug Nozzle on a J-85 Afterburning Turbojet Engine

Plug nozzle film cooling data were obtained downstream of a slot located at 42 percent of the total plug length on a J-85 engine. Film cooling reduced the aft end wall temperature as much as 150 K, reduced total pressure loss in the upstream convection cooling passages by 50 percent, and reduced estimated compressor bleed flow requirement by 14 percent compared to an all convectively cooled nozzle. Shock waves along the plug surface strongly influenced temperature distributions on both convection and film cooled portions. The effect was most severe at nozzle pressure ratios below 10 where adverse pressure gradients were most severe.

Nosek, S. M.↗

NASA NTRS1987

Turbine airfoil film cooling

The experimental data obtained in this program gives insight into the physical phenomena that occur on a film cooled airfoil, and should provide a relevant data base for verification of new design tools. Results indicate that the downstream film cooling process is a complex function of the thermal dilution and turbulence augmentation parameters with trends actually reversing as blowing strength and coolant-to-gas temperature ratio varied. The pressure surface of the airfoil is shown to exhibit a considerably higher degree of sensitivity to changes in the film cooling parameters and, consequently, should prove to be more of a challenge than the suction surface in accurately predicting heat transfer levels with downsteam film cooling.

Hylton, L. D.↗

NASA NTRS1979

Effect of shocks on film cooling of a full scale turbojet exhaust nozzle having an external expansion surface

Experimental film cooling data obtained during exploratory testing with an axisymmetric plug nozzle having external expansion and installed on an afterburning turbojet engine in an altitude test facility is presented. The shocks and local hot gas stream conditions had a marked effect on film cooling effectiveness. An existing film cooling correlation was adequate at some operating conditions but inadequate at other conditions such as in separated flow regions resulting from shock boundary layer interactions.

Straight, D. M.↗

NASA NTRS2015

Analysis of Turbine Blade Relative Cooling Flow Factor Used in the Subroutine Coolit Based on Film Cooling Correlations

Heat transfer correlations of data on flat plates are used to explore the parameters in the Coolit program used for calculating the quantity of cooling air for controlling turbine blade temperature. Correlations for both convection and film cooling are explored for their relevance to predicting blade temperature as a function of a total cooling flow which is split between external film and internal convection flows. Similar trends to those in Coolit are predicted as a function of the percent of the total cooling flow that is in the film. The exceptions are that no film or 100 percent convection is predicted to not be able to control blade temperature, while leaving less than 25 percent of the cooling flow in the convection path results in nearing a limit on convection cooling as predicted by a thermal effectiveness parameter not presently used in Coolit.

Air cooling↗

NASA NTRS1990

A computational design tool for scramjet combustor film cooling and fuel mixing predictions

A methodology for the analysis of a film cooled combustor of a hydrogen-fueled, supersonic combustion ramjet (scramjet) is presented. The SHIP3D parabolic flow code, based on an extension of the pressure-correction method of Patankar and Spalding for compressible flow is used to compute the flow field in the combustor. The effect of film cooling on the overall vehicle performance is then calculated using the one-dimensional, multistep cycle code, CPIPE. Results are presented for a conceptual combustor at Mach 16 flight conditions. It is seen that film cooling can have a significant effect on the engine performance.

Kamath, Pradeep S.↗

NASA NTRS1992

Analysis of film cooling in rocket nozzles

Progress during the reporting period is summarized. Analysis of film cooling in rocket nozzles by computational fluid dynamics (CFD) computer codes is desirable for two reasons. First, it allows prediction of resulting flow fields within the rocket nozzle, in particular the interaction of the coolant boundary layer with the main flow. This facilitates evaluation of potential cooling configurations with regard to total thrust, etc., before construction and testing of any prototype. Secondly, CFD simulation of film cooling allows for assessment of the effectiveness of the proposed cooling in limiting nozzle wall temperature rises. This latter objective is the focus of the current work. The desired objective is to use the Finite Difference Navier Stokes (FDNS) code to predict wall heat fluxes or wall temperatures in rocket nozzles. As prior work has revealed that the FDNS code is deficient in the thermal modeling of boundary conditions, the first step is to correct these deficiencies in the FDNS code. Next, these changes must be tested against available data. Finally, the code will be used to model film cooling of a particular rocket nozzle. The third task of this research, using the modified code to compute the flow of hot gases through a nozzle, is described.

Woodbury, Keith A.↗

NASA NTRS1996

The Effect of Wake Passing on Turbine Blade Film Cooling

The effect of upstream blade row wake passing on the showerhead film cooling performance of a downstream turbine blade has been investigated through a combination of experimental and computational studies. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of showerhead film hole pattern. Film effectiveness and Nusselt number values were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers. Results indicated a demonstratable reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated well by a chordwise-constant decrement of 0.094-St. Measurable spanwise film effectiveness variations were found near the showerhead region, but meaningful unsteady variations and downstream spanwise variations were not found. Nusselt numbers were less sensitive to wake and injection changes. Computations were performed using a three-dimensional turbulent Navier-Stokes code which was modified to model wake passing and film cooling. Unsteady computations were found to agree well with steady computations provided the proper time-average blowing ratio and pressure/suction surface flow split are matched. The remaining differences were isolated to be due to the enhanced mixing in the unsteady solution caused by the wake sweeping normally on the pressure surface. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

Heidmann, James David↗