Film cooling on a convex surface: influence of external pressure gradient and Mach number on film cooling performance

 The film cooling performance on a convex surface subjected to zero and favourable pressure gradient free-stream flow was investigated. Adiabatic film cooling effectiveness values were obtained for five different injection geometries, three with cylindrical holes and two with shaped holes. Heat transfer coefficients were derived for selected injection configurations. CO₂ was used as coolant to simulate density ratios between coolant and free-stream close to gas turbine engine conditions. The film cooling effectiveness results indicate a strong dependency on the free-stream Mach number level. Results obtained at the higher free-stream Mach number show for cylindrical holes generally and for shaped holes at moderate blowing rates significant higher film cooling effectiveness values compared to the lower free-stream Mach number data. Free-stream acceleration generally reduced adiabatic film cooling effectiveness relative to constant free-stream flow conditions. The different free-stream conditions investigated indicate no significant effects on the corresponding heat transfer increase due to film injection. The determined heat flux ratios or film cooling performance indicated that coolant injection with shaped film cooling holes is much more efficient than with cylindrical holes especially at higher blowing rates. Heat flux penalties can occur at high blowing rates when using cylindrical holes. Received on 29 May 2000     

Convective heat and mass transfer from a falling film to a laminar gas stream

A numerical study has been made of convective heat and mass transfer from a falling film to a laminar gas stream between vertical parallel plates. The effects of gas-liquid phase coupling, variable thermophysical properties, and film vaporization have been considered. Simultaneous mass, momentum and heat transfer between liquid film and gas stream is numerically studied by solving the respective governing equations for the liquid film and gas stream together. The influences of the inlet liquid temperature and liquid flowrate on the cooling of liquid film are examined for air-water and air-ethanol systems. Results show that the heat transfer from the gas-liquid interface to the gas stream is predominantly determined by the latent heat transfer connected with film evaporation. Additionally, better liquid film cooling is noticed for the system having a higher inlet liquid temperature or a lower liquid flowrate.     

Improvement of film cooling effectiveness with a small downstream block body

The aim of this study is to predict the improvement in film cooling performance over a flat plate through a single row of cylindrical holes with different streamwise angles by using the Ansys CFX software package. In order to improve the film cooling effectiveness, a short crescent-shaped block is placed downstream of a cylindrical cooling hole. The numerical results of the cylindrical hole without the downstream short crescent-shaped block are compared with experimental data.     

Film cooling and heat transfer with air injection through two rows of compound angle holes.

This paper describes the results of an experimental investigation into the film cooling effectiveness and the heat transfer characteristics of two staggered rows of compound angle holes. The effects of hole spacings and turbulence intensity on film cooling and heat transfer characteristic are investigated for three blowing rates; 0.5, 1.0 and 1.7. An attempt has been made to correlate the film cooling effectiveness results using a two dimensional correlation group. The increase of spanwise hole spacing results in a reduction in the film cooling effectiveness and an increase in the Stanton number. Increasing the freestream turbulence intensity has caused a significant reduction in the local film cooling effectiveness but increased the Stanton number, especially at blowing rate of 0.5.     

High resolution PIV measurements around a model turbine blade trailing edge film-cooling breakout

Film cooling downstream of a model turbine blade trailing edge has been studied experimentally. High resolution particle image velocimetry was used to obtain spatially resolved mean velocity and turbulence measurements in the immediate vicinity of the trailing edge breakout. The mean velocity measurements imply the presence of a pair of counter-rotating longitudinal vortices shed from the sides of the breakout lands. The turbulent shear stress measurements above the breakout are significantly intensified as blowing ratio is increased. These results suggest that there is a strong mixing between the film cooling slot jets and the mainstream flow which degrades the film cooling effectiveness.     

Destabilization of film boiling by means of a thermal resistance

Described is the type of vaporisation which takes place when a thermal resistance, consisting in a film of a substance of low heat conductivity, is placed between the surface of a quenched sample and the cooling liquid. This type of vaporisation, larvate boiling, is characterised by an alternate wetting/non-wetting of the solid surface.Two conditions are necessary for larvate boiling: thermal resistance and surface effusivity.Substituting larvate boiling for film boiling allows the heat flux between a solid surface at high temperature and the cooling liquid to be greatly increased.     

Analysis of cooling of a conducting fluid film of non-uniform thickness on a rotating disk

The unsteady thin conducting liquid film of non-uniform thickness on a rotating disk which is cooled axisymmetrically from below has been analysed numerically by solving the evolution equation of the free surface. Transient film profiles for different initial liquid film distributions have been obtained. The results reveal that the thinning process and film planarisation are markedly influenced by the heat dissipating or cooling parameter β, Prandtl number σ and Reynolds number Re.     

Experimental study of rotation effect on film cooling over the flat wall with a single hole

A new rotating test rig was set up to investigate the rotation effect on the film cooling over the flat wall. A simple flat blade with an inclined 30° film hole, which is parallel to the hot mainstream, was installed. And different rotation orientations were selected to simulate the blade pressure or suction side of a turbine blade. A steady liquid crystal technique was applied to obtain detailed distribution of the temperature over the blade surface. And the average adiabatic film cooling effectiveness of the area adjacent to the film hole was selected to evaluate the cooling effect. Five different rotational speeds, i.e., 0, 300, 500, 800, 1000 r/min, were considered. Experimental results indicate that the film trajectory could bend under the rotating condition. With the increase of the rotational speed, on the pressure side, the film trajectory inclines centripetally firstly and then centrifugally; whereas, on the suction side the film trajectory bends centrifugally. On the other hand, as the rotational speed increases, the cooling effect is improved firstly and then worsened when Ω > 500–600 r/min on the pressure side. On the suction side, however, the cooling effect is not sensitive to the rotational speed.     

The stability and development of tip and root vortices behind a model wind turbine

An inclined turbulent jet discharging a passive scalar into a turbulent crossflow is investigated under conditions of favorable, zero and adverse streamwise pressure gradient. Experiments are conducted in water by means of magnetic resonance velocimetry and magnetic resonance concentration measurements. The velocity ratio and density ratio are equal to one for all cases. The flow configuration is relevant to film cooling technology, the molecular properties of the fluid being immaterial in the fully turbulent regime. Under favorable pressure gradient (FPG), the streamwise acceleration tilts the jet trajectory toward the wall, which would be beneficial for the film cooling performance. However, the counter-rotating vortex pair is strengthened in the accelerating flow by streamwise stretching. Also, the crossflow boundary layer is significantly thickened by increasingly adverse pressure gradient, which affects the mass transfer from the jet. Overall, the more intense counter-rotating vortices and the thinner boundary layer associated with increasingly FPG enhance the scalar dispersion into the main flow, hampering the film cooling performance in terms of surface effectiveness.     

高超声速边界层液膜演化过程和冷却机理研究

高超声速液膜冷却技术是通过一系列狭缝或孔洞压出冷却工质,在飞行器表面边界层形成一层低温冷却膜,阻止高超声速气流对飞行器的气动加热.其作为一种主动冷却方式在高超声速飞行器表面热防护有着巨大的应用潜力.文章采用数值方法,结合VOF模型,研究25 km飞行高度和Ma=5气流条件下的液膜铺展情况,并通过不同冷却工质的入射速度、角度、表面张力和黏性系数条件,讨论了液膜在平板上的演化过程和冷却机理.结果表明,在气流作用下,液膜向壁面下游发展,液膜的存在导致边界层分离,连续液膜会在一定位置断裂为液块,然后进一步破碎为液滴.入射条件和液体性质的改变,会影响液膜沿流向的发展,具体表现在连续液膜断裂点的位置和连续液膜的厚度.在所设定的计算域内,壁面热流降低了80%～95%,液膜对壁面的冷却效率随着液膜形态的变化而变化.     

Numerical simulation of the effect of plasma aerodynamic actuation on improving film hole cooling performance

The primary goal of this paper is to study film cooling performance for a cylindrical hole with plasma aerodynamic actuation. The simulation model of plasma aerodynamic actuation on improving film hole cooling effectiveness was established. The heat effect of plasma aerodynamic actuation model was taken into consideration. It was firstly found that the velocity and blowing ratio greatly affect the film cooling effectiveness. Then, position, power input, and the number of plasma actuators were particularly investigated.     

Film cooling effectiveness from trenched shaped and compound holes

This paper presents a comparative-numerical investigation on film cooling from a row of simple and compound-angle holes injected at 35° on a flat plate with four film cooling configurations: (1) cylindrical film hole; (2) 15° forward diffused film hole; (3) trenched cylindrical film hole; (4) trenched 15° forward-diffused film hole. All simulations are at fixed density ratio of 1.6, blowing ratio of 1.25, length-to-diameter L/D = 4 and pitch-to-diameter ratio of 3.0. The effect of length-to-diameter ratio on film cooling has been also investigated using L/D in the range of 1–8. Computational solutions of the steady, Reynolds-averaged Navier–Stokes equations have been obtained using a finite volume method. It has been found that the shape of the hole and the trenched holes can significantly affect the film cooling flow over the protected surface. Further, it has been shown that the film cooling effectiveness by trenched shaped holes is higher than all other configurations both in spanwise and streamwise specially downstream of the injection. Also, a trenched compound angle injection shaped hole produces much higher film cooling protection than the other configurations investigated in the present paper. The length-to-diameter ratio of trenched holes was found to have a significant effect on film cooling effectiveness and the spread of the coolant jets.     

Film cooling by injection into a turbulent boundary layer

Most papers on film cooling concern injection of a homogeneous gas. Stollery et al. [1] examined the case of tangential injection of gas into a boundary layer, the specific heat_63-01 differing little from that of the main flow,_63-02.Here we examine the effectiveness of film cooling of a thermally isolated planar wall by local supply to a turbulent boundary layer.     

Berechnung der Störung der Brennkammer-Filmkühlung durch Einblasen von Luftstrahlen mit Hilfe einer Analogiebetrachtung zwischen Strahl und Zylinder

The influence of cold air jet injection upon the film cooling of combustion chamber walls is investigated theoretically and experimentally. Considering the cold air jet acting like a cylinder in the combustion gas and film cooling streams, an equation can be derived theoretically which describes the disturbance of the film cooling effectiveness downstreams of the air jet injection. The experiments, where single cold air jets are injected normal to the film cooled test chamber wall, show considerable decreases of the film cooling effectiveness downstream of the injection points. The same effect is noted in real combustion chambers of aircraft gas turbine engines. Additional experiments with cylinders simulating the disturbance effect of the air jets prove the analogy consideration between jet and cylinder. The theoretically predicted cooling effectiveness is in agreement with own experiments as well as other test results.     

Numerical study on film cooling effectiveness from shaped and crescent holes

This paper presents a comparative numerical investigation on film cooling from a row of holes injected at 35° on a flat plate with three film cooling configurations, including cylindrical hole, 15° forward diffused shaped hole, and new crescent hole. All simulations are conducted at blowing ratio of 0.6 and 1.25, length-to-diameter ratio of four and pitch-to-diameter ratio of three. Computational solutions of the steady, Reynolds averaged Navier–Stokes equations are obtained using a finite volume method. Previous successful application of a two-layer turbulence model to cylindrical hole is extended to predict film cooling for the different hole geometries. It has been found that the film cooling effectiveness of cylindrical holes obviously declined along with increasing the blowing ratio. While the forward diffused shaped hole presents a marked improvement, with a higher effectiveness at the lateral area between adjacent holes. By comparison, the crescent hole exhibits the highest film cooling effectiveness among the three configurations both in spanwise and streamwise especially downstream of the intersection of the two holes. Also, the crescent hole can restrain the vortex intensity, and then enhance the film cooling effectiveness.     

Numerical modelling of film cooling from converging slot-hole

This paper presents a numerical prediction of a new 3D film cooling hole geometry, the converging slot-hole or console. The console geometry is designed in order to improve the heat transfer and aerodynamic loss performance of turbine vane and rotor blade cooling systems without loosing the mechanical strength of a row of discrete holes. The cross section of the console changes from a circular shape at the inlet to a slot at the exit. Previous successful application of a new anisotropic DNS based two-layer turbulence model to cylindrical and shaped hole injections is extended to predict film cooling for the new console geometry. The suitability of the proposed turbulence model for film cooling flow is validated by comparing the computed and the measured wall-temperature distributions of cylindrical hole injections. The result shows that the anisotropic eddy-viscosity/diffusivity model can correctly predict the spanwise spreading of the temperature field and reduce the strength of the secondary vortices. Comparative computations of the adiabatic film cooling effectiveness associated with the three geometries tested in the present study (cylindrical, shaped, and console) show that the new console film-cooling hole geometry is definitely superior to the other geometries as shown by the uniform lateral spreading of the effectiveness with a slight enhancement downstream of the intersection of the two consoles.     

Asymptotic model of slow three-dimensional liquid film flow in a space drop catcher

Slow steady-state film flows formed on the inner surface of a drop catcher funnel due to inertial deposition of drops of a dispersed working matter in the spacecraft cooling system are considered. A limiting asymptotic model of slow three-dimensional coolant film flow is constructed assuming that the deposited drops transfer all their mass, momentum, and energy to the film described by the equations of creeping viscous fluid flow in a thin layer of a priori unknown thickness. A first-order quasi-linear partial differential equation for the film thickness is derived. The shape of the film surface is investigated numerically as a function of parameters using the method of characteristics. The range of optimum parameters ensuring the steady-state film flow is found. The limits of existence of the solutions corresponding to the limiting model proposed are investigated.     

高超声速溢流冷却实验研究

高超声速溢流冷却是一种新型的飞行器热防护方法,基本思想为:在高热流区布置溢流孔,控制冷却液以溢流方式流出,之后通过飞行器表面摩阻作用展布为液膜,形成热缓冲层以降低飞行器表面热流. 目前,溢流冷却技术还处于探索阶段,实现工程应用前还需开展大量的实验验证和机理研究工作. 本文首次开展溢流冷却的实验研究工作,采用热流测量、液膜厚度测量及液膜流动特性观测技术,搭建了完善的溢流冷却风洞实验平台,对溢流冷却热防护性能和高超声速条件下液膜流动规律进行了初步研究. 研究表明:(1) 高超声速流场中通过溢流能够在飞行器表面形成液膜并有效隔离外部高温气流,可降低飞行器表面热流率;(2) 楔面上的液膜前缘流动是一个逐渐减速的过程,增加冷却液流量液膜厚度变化不明显,但液膜前缘运动速度增大;(3) 液膜层存在表面波,在时间和空间方向发生演化,导致液膜厚度的微弱扰动;(4) 液膜层存在横向展宽现象,即液膜层宽度大于溢流缝宽度. 原因是液膜层与流场边界层条件不匹配,存在压力梯度,迫使冷却液向低压区流动,从而展宽液膜层,并且流量越高,横向展宽现象越明显.      

Cooling of a viscoelastic film during unsteady extrusion from an annular die

The unsteady extrusion of a viscoelastic film from an annular and axisymmetric die is examined. External, elastic, viscous and inertia forces deform the film, which is simultaneously cooled via forced convection to the ambient air. This moving boundary problem is solved by mapping the liquid/air interfaces onto fixed ones and by employing a regular perturbation expansion for all the dependent variables. The ratio of the film thickness to its inner radius at the exit of the die is used as the small parameter in the perturbation expansion. The fluid mechanical aspects of the process depend on the Stokes, Deborah, Reynolds, and Capillary numbers. The heat transfer in the film and to the environment gives rise to four additional dimensionless groups: the Peclet, Biot and Brinkman numbers and the activation energy, which determines the temperature dependence of fluid viscosity and elasticity. A variable heat transfer coefficient is also considered. For typical fluid properties and process conditions, the Peclet number is very large. In this case it is the ratio of the Biot to the Peclet number, the Stanton number, which arises in the energy conservation equation. It is shown that film cooling becomes important when the Stanton number and/or the activation energy are in the high-end of their typical values. In such cases, the cooling of the parison leads to a more uniform flow and shape for the film. The influence on the process of a variable heat transfer coefficient and the Brinkman number is small. Received: 7 April 1999/Accepted: 10 August 1999     

Specific impulse losses due to friction and dispersion in a gas-film cooled liquid rocket engine nozzle

Conclusions The proposed method and program for calculating the compressible turbulent boundary layer in rocket engine nozzles with gas film cooling make it possible to determine the specific impulse losses due to friction, the heat fluxes and other characteristics of the flow. The calculations are based on the numerical solution of the equations of gas dynamics in the boundary layer approximation using a three-parameter differential turbulence model.The calculations for nozzles without film cooling showed that the contours occupying a narrow interval between the families of contours with uniform and variational characteristics have the minimum impulse losses due to friction and dispersion. In contrast to the known results, the loss minimum is displaced relative to nozzles with a variational characteristic (Rao nozzles) towards truncated nozzles with a uniform characteristic.The dependence of the maximum heat transfer to the wall in the critical throat section of the nozzle on the rate of flow of fuel into the film has been determined for nozzles with film cooling. It is shown that as the film flow rate increases, the friction losses decrease, and the minimum of the impulse losses due to friction and dispersion is shifted towards the contours with a variational characteristic, which have the minimum dispersion losses. The total impulse losses, which take into account the change in the fuel component ratio in the flow core due to the diversion of part of the fuel into the film, increase with increase in the film flow rate.The results of our numerical investigation of the effect of the contour shape and film flow rate indicate that the contour with a variational characteristic, which has near-minimum specific impulse losses due to friction and dispersion, should be used as the optimum contour for LRE nozzles.Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No.4, pp. 82–93, May–June, 1993.deceased.The authors wish to thank their colleagues at the Énergomash NPO L. P. Vereshchak and L. K. Danilyuk for assisting with the calculations, the participants in G. A. Lyubimov's seminar for discussing the results obtained, and D. A. Mel'nikov, U. G. Pirumov, and A. A. Sergienko for valuable advice.