Importance of combining convection with film cooling
01 Jan 1972-
TL;DR: In this paper, the interaction of film and convection cooling and its effect on wall cooling efficiency was investigated analytically for two cooling schemes for advanced gas turbine applications, full coverage-and counterflow-film cooling.
Abstract: The interaction of film and convection cooling and its effect on wall cooling efficiency is investigated analytically for two cooling schemes for advanced gas turbine applications. The two schemes are full coverage- and counterflow-film cooling. In full coverage film cooling, the cooling air issues from a large number of small discrete holes in the surface. Counterflow film cooling is a film-convection scheme with film injection from a slot geometry. The results indicate that it is beneficial to utilize as much of the cooling air heat sink as possible for convection cooling prior to ejecting it as a film.
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TL;DR: In this article, the authors developed a test facility for investigating full coverage discrete hole wall cooling for gas turbine combustion chamber wall cooling, and a low temperature test condition of 750k mainstream temperature and 300k coolant temperature was used to investigate the influence of coolant flow rate at a constant cross flow Mach number.
Abstract: The development of a test facility for investigating full coverage discrete hole wall cooling for gas turbine combustion chamber wall cooling is described. A low temperature test condition of 750K mainstream temperature and 300K coolant temperature was used to investigate the influence of coolant flow rate at a constant cross flow Mach number. Practical combustion conditions of 2100K combustor temperature and 700K coolant temperature are investigated to establish the validity of applying the low temperature results to practical conditions. For both situations a heat balance programme, taking into account the heat transfer within the wall was used to compute the film heat transfer coefficients. The mixing of the coolant air with the mainstream gases was studied through boundary layer temperature and CO2 profiles. It was shown that entrainment of hot flame gases between the injection holes resulted in a very low 'adiabatic' film cooling effectiveness.
20 citations
TL;DR: This article reviewed some of the work on advanced liner cooling techniques, specifici~lly laminated porous wall cooling, angled-multihole (effusion) cooling ant1 composite metal matrix liner cooling.
Abstract: This paper briefly reviews some of the work on advanced liner cooling techniques - specifici~lly laminated porous wall cooling, angled-multihole (effusion) cooling ant1 composite metal matrix liner cooling. The concept definition, heat transfer design procedure and design problems including key materials and fabrication considerations associated with each basic concept will be reviewed. Published rig and engine experience of aircraft engine manufacturers and reseirch organizations will be cited. Some logical extensions of the current liner cooling schemes are suggested for future applications.
17 citations
01 Feb 1976
TL;DR: Aerodynamic effects of trailing edge geometry, hole size, angle, spacing, and shape were studied in two-and three-dimensional cascades and in a warm turbine test series as discussed by the authors.
Abstract: Aerodynamic effects of trailing edge geometry, hole size, angle, spacing, and shape were studied in two- and three-dimensional cascades and in a warm turbine test series. Heat transfer studies were carried out in various two- and three-dimensional test facilities in order to provide corresponding heat transfer data. Results are shown in terms of cooling effectiveness and aerodynamic efficiency for various coolant fractions, coolant-primary temperature ratios, and cooling configurations.
7 citations
01 Jul 1978
TL;DR: In this paper, a computer program that calculates the coolant flow and the metal temperatures of a full-coverage-film-cooled vane or blade was developed, based on compressible, one-dimensional fluid flow and on one dimensional heat transfer.
Abstract: A computer program that calculates the coolant flow and the metal temperatures of a full-coverage-film-cooled vane or blade was developed. The analysis was based on compressible, one-dimensional fluid flow and on one-dimensional heat transfer and treats the vane or blade shell as a porous wall. The calculated temperatures are average values for the shell outer-surface area associated with each film-cooling hole row. A thermal-barrier coating may be specified on the shell outer surface, and centrifugal effects can be included for blade calculations. The program is written in FORTRAN 4 and is operational on a UNIVAC 1100/42 computer. The method of analysis, the program input, the program output, and two sample problems are provided.
5 citations
01 Feb 1980
TL;DR: The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally as mentioned in this paper, and the results showed that metal temperatures in the leading edge region of the vanes were significantly lower than those of the non-coated vanes.
Abstract: The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally. Current and advanced gas turbine engine conditions were simulated at reduced temperatures and pressures. Airfoil metal temperatures were significantly reduced, both locally and on the average, by use of the the coating. However, at low gas Reynolds number, the ceramic coating tripped a laminar boundary layer on the suction surface, and the resulting higher heat flux increased the metal temperatures. Simulated coating loss was also investigated and shown to increase local metal temperatures. However, the metal temperatures in the leading edge region remained below those of the uncoated vane tested at similar conditions. Metal temperatures in the trailing edge region exceeded those of the uncoated vane.
2 citations
References
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TL;DR: In this paper, basic heat transfer and flow friction design data are presented for nine offset rectangular plate-fin surfaces, seven of the test cores were constructed of aluminum and two of stainless steel.
Abstract: : Basic heat transfer and flow friction design data are presented for nine offset rectangular plate-fin surfaces. Seven of the test cores were constructed of aluminum and two were constructed of stainless steel. One of the configurations had a heat transfer area density ratio of 1772 sq ft/cu ft, one of the most compact plate-fin surface ever tested at Stanford University. The present results are compared with the offset rectangular plate-fin surfaces previously tested at Stanford. Comparisons are also provided with a louvered fin and a perforated fin plate-fin surface, in terms of heat transfer power versus flow friction power and flow area characteristics. The various kinds of mal-flow distribution through the test cores and their effects on the test results for heat transfer and flow friction are examined. (Author)
93 citations