Mainstream Mass Flow Ratio Research Articles

Effects of casing air injection on film cooling and aerodynamic performances of an axial turbine cascade

Rotor casing injection is one of the technical means for tip leakage flow control and casing thermal management. In the current investigation, a series of numerical simulations with the unsteady sliding mesh method are conducted to illustrate the effects of injection location, tip clearance, and shaped injection hole on the aerodynamic and cooling performance of casing air injection in an axial turbine cascade, under three injection mass flow ratios (0.15%, 0.30%, and 0.45%). The results indicate that the casing injection at X/Ca = 0.1 performs the best in reducing the tip leakage mass flow rate, but the highest isentropic efficiency appears when the injection holes are located at X/Ca = 0.3. The increase of the tip clearance is detrimental to the aerodynamic and film cooling performances of the casing injection scheme. Moreover, the application of shaped holes (converging slot hole and fan-shaped hole) plays a positive role in improving film cooling effectiveness and increasing isentropic efficiency. Especially under a large coolant to mainstream mass flow ratio of 0.45%, the results show that the fan-shaped hole improves the spatially-averaged casing film cooling effectiveness by approximately 16% relative to the cylindrical hole.

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.

Influence of coolant mass flow rate on the endwall conjugate heat transfer

The switch from diffusive combustion to premixed combustion in a modern gas turbine will change the combustor exit temperature profile to a more uniform one. As a result, the temperature near the inner and outer shroud of the turbine vane will increase which is a big challenge for the endwall cooling. Two typical vane endwall configurations with different cooling arrangements are investigated both experimentally and numerically in this study. One endwall has film cooling method only while the other has both film cooling and impingement cooling. The coolant to mainstream mass flow ratio (MFR) changes from 0.8% to 2.5% in the numerical investigation. The numerical method is validated by the experimental data. Detailed analysis of the coolant mass flow rate effect on the endwall conjugate heat transfer is conducted based on the numerical results. The results indicate that the speed of the overall effectiveness increment slows down when the MFR increases. As the MFR increases, the advantage of the combination of film and impingement enlarges. The warming effect of the coolant increases as the MFR increases which leads to a lower coolant discharge temperature.

Heat Transfer Performance of a Transonic Turbine Blade Passage in the Presence of Leakage Flow Through Upstream Slot and Mateface Gap With Endwall Contouring

Comparison of heat transfer performance of a nonaxisymmetric contoured endwall to a planar baseline endwall in the presence of leakage flow through stator–rotor rim seal interface and mateface gap is reported in this paper. Heat transfer experiments were performed on a high turning turbine airfoil passage at Virginia Tech's transonic blow down cascade facility under design conditions for two leakage flow configurations—(1) mateface blowing only, (2) simultaneous coolant injection from the upstream slot and mateface gap. Coolant to mainstream mass flow ratios (MFRs) were 0.35% for mateface blowing only, whereas for combination blowing, a 1.0% MFR was chosen from upstream slot and 0.35% MFR from mateface. A common source of coolant supply to the upstream slot and mateface plenum made sure the coolant temperatures were identical at both upstream slot and mateface gap at the injection location. The contoured endwall geometry was generated to minimize secondary aerodynamic losses. Transient infrared thermography technique was used to measure endwall surface temperature and a linear regression method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness, assuming a one-dimensional (1D) semi-infinite transient conduction. Results indicate reduction in local hot spot regions near suction side as well as area averaged HTC using the contoured endwall compared to baseline endwall for all coolant blowing cases. Contoured geometry also shows better coolant coverage further along the passage. Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed.

Experimental Investigation of Endwall Heat Transfer With Film and Impingement Cooling

The switch from diffusive combustion to premixed combustion in a modern gas turbine changes the combustor exit temperature profile to a more uniform one. This will directly affect the cooling of the first stage vane especially the endwall region. A typical endwall configuration with matched nondimensional parameters to the engine condition was investigated experimentally in this study. Two endwall cooling arrangements at four different coolant to mainstream mass flow ratios (MFR) were tested in a linear cascade. Detailed measurements of pressure distribution, heat transfer coefficient, adiabatic film cooling effectiveness, and overall effectiveness of the endwall were performed. The temperature-sensitive paint (TSP) and pressure-sensitive paint (PSP) were used to acquire these parameters. The conjugate heat transfer characteristic of endwall with film cooling and impingement cooling was discussed. Moreover, the influence of coolant mass flow rate on conjugate heat transfer of endwall was analyzed. One- and two-dimensional methods for overall effectiveness prediction based on experimental data for separate parameters and correlations were also studied.

Effects of injection conditions and Mach number on unsteadiness arising within coolant jets over a pressure side vane surface

The thermal performance of a gas turbine airfoil with a cooled pressure side and a trailing edge cutback is investigated and discussed in relationship with the unsteady behavior of coolant injection. The focus is on the pressure side coolant injection through discrete holes. The cascade was tested at an exit Mach number of 0.2 and 0.6 for different coolant to mainstream mass flow ratios. Laser Doppler Velocimetry (LDV) and high speed flow visualizations were used to investigate the unsteady mixing process taking place between coolant and main flow downstream of the holes. This behavior was correlated with the film cooling effectiveness distributions over the vane surface. In the first row, hairpin clockwise rotating vortices and counterclockwise shear layer vortices were observed, depending on the coolant to mainstream velocity ratio. The second row turned out to be characterized by a lower velocity ratio and lower turbulent activity, consistently with the higher thermal protection observed also at high injection rates. The increase in mainstream Mach number kept the jets closer to the wall, resulting in lower turbulent characteristics and lower mixing, consistently with higher thermal protection.

Experimental investigation of the effects of blowing conditions and Mach number on the unsteady behavior of coolant ejection through a trailing edge cutback

The present paper shows the results of an experimental investigation into the unsteadiness of coolant ejection at the trailing edge of a highly loaded nozzle vane cascade. The trailing edge cooling scheme features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is also ejected through two rows of cylindrical holes placed upstream of the cutback. Tests were performed with a low inlet turbulence intensity level (Tu1=1.6%), changing the cascade operating conditions from low speed (M2is=0.2) up to high subsonic regime (M2is=0.6), and with coolant to main stream mass flow ratio varied within the 0.5–2.0% range. Particle Image Velocimetry (PIV) and flow visualizations were used to investigate the unsteady mixing process taking place between coolant and main flow downstream of the cutback, up to the trailing edge. For all the tested conditions, the results show the presence of large coherent structures, which presence is still evident up to the trailing edge. Their shape and direction of rotation change with injection conditions, as a function of coolant to mainstream velocity ratio, strongly influencing the thermal protection capability of the injected coolant flow. The Mach number increase is only responsible for a positioning of such vortical structures closer to the wall, while the Strouhal number almost remains unchanged.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Film cooling of a contoured endwall nozzle vane through fan-shaped holes

The present paper shows the results of an experimental investigation into a nozzle vane cascade with endwall contouring and film cooling. In a film cooled endwall the front area, especially close to the leading edge, is one of the most critical regions to be cooled; in case of a contoured endwall the thermal protection becomes even more difficult because of the endwall curvature effect. Two film cooling hole geometries with four rows of holes were tested: cylindrical holes and conical expanded holes. Tests were performed at low speed ( M 2 is = 0.2) and low inlet turbulence intensity level, with coolant to main stream mass flow ratio varied within the 0.5–2.5% range. Aerodynamic and thermal performances were evaluated in every injection condition. Compared to cylindrical holes, results showed that shaped film cooling holes generate slightly higher secondary losses and provide a better thermal coverage, at the expense of a higher coolant flow injection.

Endwall Film Cooling Effects on Secondary Flows in a Contoured Endwall Nozzle Vane

The present paper investigates the effects of endwall injection of cooling flow on the aerodynamic performance of a nozzle vane cascade with endwall contouring. Tests have been performed on a seven vane cascade with a geometry typical of a real gas turbine nozzle vane. The cooling scheme consists of four rows of cylindrical holes. Tests have been carried out at low speed (Ma2is=0.2) with a low inlet turbulence intensity level (1.0%) and with a coolant to mainstream mass flow ratio varied in the range from 0% (solid endwall) to 2.5%. Energy loss coefficient, secondary vorticity, and outlet angle distributions were computed from five-hole probe measured data. Contoured endwall results, with and without film cooling, were compared with planar endwall data. Endwall contouring was responsible for a significant overall loss decrease, as a result of the reduction in both profile and planar side secondary flows losses; a loss increase on the contoured side was instead observed. Like as for the planar endwall, even for the contoured endwall, coolant injection modifies secondary flows, reducing their intensity, but the relevance of the changes is reduced. Nevertheless, for all the tested injection conditions, secondary losses on the contoured side are always higher than in the planar case, while contoured cascade overall losses are lower. A unique minimum overall loss injection condition was found for both tested geometries, which corresponds to an injected mass flow rate of about 1.0%.

Film-Cooling Effectiveness on a Rotating Turbine Platform Using Pressure Sensitive Paint Technique

Film-cooling effectiveness is measured on a rotating turbine blade platform for coolant injection through discrete holes using pressure sensitive paint technique. Most of the existing literatures provide information only for stationary endwalls. The effects of rotation on the platform film-cooling effectiveness are not well documented. Hence, the existing three-stage turbine research facility at the Turbomachinery and Flow Performance Laboratory, Texas A&M University was redesigned and installed to enable coolant gas injection on the first stage rotor platform. Two distinct coolant supply loops were incorporated into the rotor to facilitate separate feeds for upstream cooling using stator-rotor gap purge flow and downstream discrete-hole film cooling. As a continuation of the previously published work involving stator-rotor gap purge cooling, this study investigates film-cooling effectiveness on the first stage rotor platform due to coolant gas injection through nine discrete holes located downstream within the passage region. Film-cooling effectiveness is measured for turbine rotor frequencies of 2400 rpm, 2550 rpm, and 3000 rpm corresponding to rotation numbers of Ro=0.18, 0.19, and 0.23, respectively. For each of the turbine rotational frequencies, film-cooling effectiveness is determined for average film-hole blowing ratios of Mholes=0.5, 0.75, 1.0, 1.25, 1.5, and 2.0. To provide a complete picture of hub cooling under rotating conditions, simultaneous injection of coolant gas through upstream stator-rotor purge gap and downstream discrete film-hole is also studied. The combined tests are conducted for gap purge flow corresponding to coolant to mainstream mass flow ratio of MFR=1% with three downstream film-hole blowing ratios of Mholes=0.75, 1.0, and 1.25 for each of the three turbine speeds. The results for combined upstream stator-rotor gap purge flow and downstream discrete holes provide information about the optimum purge flow coolant mass, average coolant hole blowing ratios for each rotational speed, and coolant injection location along the passage to obtain efficient platform film cooling.

End-Wall Film Cooling Through Fan-Shaped Holes With Different Area Ratios

The present paper reports on the aerothermal performance of a nozzle vane cascade, with film-cooled end walls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two end-wall geometries with different area ratios have been compared. Tests have been carried out at low speed (M=0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through three-dimensional (3D) aerodynamic measurements, by means of a miniaturized five-hole probe. Adiabatic effectiveness distributions have been determined by using the wide-band thermochromic liquid crystals technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performances of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

Effect of coolant temperature and mass flow on film cooling of turbine blades

A three-dimensional Navier-Stokes code has been used to study the effect of coolant temperature, and coolant to mainstream mass flow ratio on the adiabatic effectiveness of a film-cooled turbine blade. The blade chosen is the VKI rotor with six rows of cooling holes including three rows on the shower head. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900K and stagnation pressure = 3 MPa. Generally, the adiabatic effectiveness is lower for a higher coolant temperature due to nonlinear effects via the compressibility of air. However, over the suction side of shower-head holes, the effectiveness is higher for a higher coolant temperature than that for a lower coolant temperature when the coolant to mainstream mass flow ratio is 5% or more. For a fixed coolant temperature, the effectiveness passes through a minima on the suction side of shower-head holes as the coolant to mainstream mass flow ratio increases, while on the pressure side of shower-head holes, the effectiveness decreases with increase in coolant mass flow due to coolant jet lift-off. In all cases, the adiabatic effectiveness is highly three-dimensional.

LEADING EDGE FILM-COOLING EFFECTS ON TURBINE BLADE HEAT TRANSFER

A three-dimensional Navier-Stokes code has been used to study the effect ofspanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature 1900 K and stagnation pressure 3 MPa. It is found that with the ratio of coolant to mainstream mass flow being fixed, reducing the spanwise pitch for shower-head holes (P) from 7.5d to 3.0d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P / d = 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface owing to coolant jet lift-off. For P / d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant-to-mainstream-mass-flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dim...