Heat transfer measurements under steady, quasi-steady and unsteady flow conditions are discussed. In continuing measurements on the influence of a plane steady wake flow as well as on the effects of grid-produced free-stream turbulence, the present paper describes the effects of the superposition of the free-stream turbulence with the wake flow. Of special interest is a comparison of the wake produced by a leading airfoil with that of a cylindrical bar in cross flow. The similarity of the wakes as far as macrostructure of the turbulence and velocity is concerned is documented using LDV and heat transfer measurements. In extending the modeling concept, a rotating wake generator is employed simulating the wake of a blade with cylindrical bars in cross flow. The comparison of the theoretical and experimental results is presented.

The flow phenomena of wakes shed by upstream blade rows is a well–known problem in turbomachinery which influences blade forces, vibrations, losses and heat transfer. With respect to the heat load to turbine blades this problem becomes even more complex because of the interaction between wake, potential flow and the boundary layer along the surface of the airfoil. Experimentally evaluated mean heat-transfer coefficients obtained under different unsteady initial conditions are reported. The heat–transfer measurements have been carried out in the cascade test facility at the ITS in Karlsruhe using a rotating bar wake generator placed upstream of the cascade to simulate the wake passing process. The variation of the wake parameters includes different wake passing frequencies, cascade inlet Reynolds numbers and wake inclination angles. In addition, the relevant parameters of the unsteady wake have been measured by means of a fixed hot–wire anemometer using the ensemble average technique. The results are compared to those from the literature for the wake of a cylinder in cross–flow. They also serve as experimental base for parallel theoretical analyses.

At the Institute for Thermal Turbomachinery, University of Karlsruhe (ITS), detailed theoretical and experimental studies concerning ceramic gas turbine components are performed. For the analysis of the reliability of ceramic components by finite element calculations the numerical code CERITS has been developed which follows a strategy similar to that suggested by Gyekenyesi. CERITS determines the failiure probability of ceramic components with reference to volumetric flaws. Different fracture criteria can be considered. CERITS is coupled to the finite element code ADINA which facilitates the prediction of the required temperature and stress distributions. The procedure is demonstrated utilizing gas turbine combustor components.

At the Institute for Thermal Turbomachinery, University of Karlsruhe (ITS), theoretical and experimental investigations on ceramic gas turbine components are performed. For the reliability analysis by finite element calculations the computer code CERITS has been developed. This code is used to determine the fast fracture reliability of ceramic components subjected to polyaxial stress states with reference to volumetric flaws and was presented at the 1990 IGTI Gas Turbine Conference. CERITS-L now includes subcritical crack growth. With the new code CERITS-L, failure probabilities of ceramic components can be calculated under given load situations versus time. In comparing these time dependent failure probabilities with a given permissible failure probability, the maximum operation time of a component can be determined. The considerable influence of the subcritical crack growth upon the life time of ceramic components is demonstrated at the flame tube segments of the ITS ceramic combustor.

At the Institute for Thermal Turbomachinery in Karlsruhe a small research combustor was developed to investigate the applicability of high temperature loaded ceramic components under near ideal engine conditions. The design concept includes film evaporation of the liquid fuel and staged combustion with variable air supply. The combustor liner is segmented and consists of ceramic elements. For the prediction of the failure probability the newly developed fracture statistic code CERITS was used. The key features of the combustor and the test facility are described and preliminary experimental results of the first test phase are reported with emphasis on emission measurements and durability of the ceramic components. The determining parameters under consideration are thermal power, air/fuel ratio and the temperature of the preheated inlet air.

In continuation of earlier studies comparing heat transfer results for flow over a backward facing step and a jet in crossflow, a comprehensive data base for slot film cooling is presented. Specifically, heat transfer measurements adjacent to the injection slot for high blowing ratios were of interest. The heat transfer distributions for these conditions are of increasing significance in modern combustor liners. The new data base includes velocity and temperature profile measurements, adiabatic wall temperature and heat transfer measurements and will serve as a data base for testing newly developed numerical models.

The present study compares measured and computed heat transfer coefficients for high speed boundary layer nozzle flows under engine Reynolds-number conditions ( U ∞ = 230 ÷ 880 m / s , Re * = 0.37 ÷ 1.07 · 10 6 ). Experimental data have been obtained by heat transfer measurements in a two-dimensional, non-symmetric, convergent-divergent nozzle. The nozzle wall is convectively cooled using water passages. The coolant heat transfer data and nozzle surface temperatures are used as boundary conditions for a three-dimensional finite-element code which is employed to calculate the temperature distribution inside the nozzle wall. Heat transfer coefficients along the hot gas nozzle wall are derived from the temperature gradients normal to the surface. The results are compared with numerical heat transfer predictions using the low Reynolds-number k -ε turbulence model by Lam and Bremhorst. Influence of compressibility in the transport equations for the turbulence properties is taken into account by using the local averaged density. The results confirm that this simplification leads to good results for transonic and low supersonic flows.

New gas turbine combustor designs are developed to reduce pollutant and NO x -emissions. In these new combustors, the formation of carbonaceous deposits, especially in prevaporizers, affects the reliablility and effectiveness of operation. To avoid deposits, a detailed knowledge of the origins and mechanisms of formation is required. To obtain a deeper insight, the phenomena were studied systematically. The deposits under consideration show differing characteristics suggesting more than one formation mechanism in the combustor. Consequently, the primary goal was to identify the formation mechanisms and, subsequently, to simulate the mechanisms under well-defined conditions in bench tests for determining the relevant parameters of deposit build-up. The mechanisms of formation were identified based on the properties of the deposits in the combustion chamber. In order to characterize the deposits, physical and chemical analysis techniques were utilized. In summary, tests and numerical predictions identified two major paths of formation: a deposit build-up resulting from flame products such as soot or coked droplets and a deposit build-up resulting from liquid fuel impinging the wall accompanied with chemical reactions at the wall. The deposits caused by fuel droplet impingement were intensively studied in bench tests. In analyzing the processes, the influence of wall temperature, fuel composition, and the oxygen content in the environment is shown in detail. In addition, the importance of thermal instabilities of the fuel, previously studied under fuel supply system conditions, is demonstrated for a deposit formation inside a combustion chamber.

A theory is presented for calculating the fluctuations in a laminar boundary layer when the free stream is turbulent. The kinetic energy equation for these fluctuations is derived and a new mechanism is revealed for their production. A methodology is presented for solving the equation using standard boundary layer computer codes. Solutions of the equation show that the fluctuations grow at first almost linearly with distance and then more slowly as viscous dissipation becomes important. Comparisons of calculated growth rates and kinetic energy profiles with data show good agreement. In addition, a hypothesis is advanced for the effective forcing frequency and free-stream turbulence level which produce these fluctuations. Finally, a method to calculate the onset of transition is examined and the results compared to data.

The emphasis of the present study is to understand the effects of various flowfield and geometrical parameters in the nearfield region of a scaled-up film-cooling hole on a flat test plate. The effect of these different parameters on adiabatic wall effectivenesses, heat transfer coefficients, discharge coefficients and the near-hole velocity field will be addressed. The geometrical parameters of concern include several angles of inclination and rotation of a cylindrical film-cooling hole and two different hole shapes — a fanshaped hole and a laidback fanshaped hole. The fluid dynamic parameters include both the internal and external Mach number as well as the mainstream-to-coolant ratios of total temperature, velocity, mass flux, and momentum flux. In particular, the interaction of a film-cooling jet being injected into a transonic mainstream will be studied. This paper includes a detailed description of the test rig design as well as the measuring techniques. Firstly, tests revealing the operability of the test rig will be discussed. Finally, an outlook of the comprehensive experimental and numerical program will be given.

At the Institut für Thermische Strömungsmaschinen in Karlsruhe new design concepts for thermally high-loaded ceramic gas turbine components were developed. The present concept is based on a load-oriented segmentation in combination with a flexible suspension and thermal insulation of the ceramic structure. The concept was applied to a flame tube for a small gas turbine. In order to ensure real operating conditions, a Klöckner Humboldt Deutz T216 type gas turbine was used as test bed for the ceramic combustor. The paper gives a description of the combustor and the test rig. Furthermore, experimental results of the engine tests with special emphasis on the liner wall temperature distribution for various steady and transient operating conditions are presented. A major result of the tests is that the design concept proved to be reliable under real engine conditions. After more than 100 hours no failure of the ceramic parts occured. In order to determine the thermal load of the ceramic flame tube under real conditions, the experimental investigations are supported by numerical calculations.

At the Institute for Thermal Turbomachinery, University of Karlsruhe (ITS) a new methodology for designing thermally loaded ceramic components based on numerical analyses of the thermal and structural behaviour is developed. Major issues are an effective optimization of the geometry and an assessment of the load. To reduce thermally induced stresses in shell structures exposed to hot gas flow a precise description of the major influencing parameters is required. Based on these relations a thermal optimization of the component is possible. Moreover, the compensation of thermal strains and the adjustment of local stiffnesses for reducing the tensile stresses and the failure probability has to be considered. In designing a first stage vane of a stationary gas turbine the proposed guidelines are verified. The advantages of a systematic approach are demonstrated. The multi-purpose numerical procedures using the guidelines allow an adjustment of the shape to thermally and mechanically induced loads and offer new possibilities to meet the specific demands of ceramic materials.

One viable option to improve cooling methods used for gas turbine blades is to optimize the geometry of the film-cooling hole. To optimize that geometry, effects of the hole geometry on the complex jet-in-crossflow interaction need to be understood. This paper presents a comparison of detailed flowfield measurements for three different single, scaled-up, hole geometries all at a blowing ratio and density ratio of unity. The hole geometries include a round hole, a hole with a laterally expanded exit, and a hole with a forward-laterally expanded exit. In addition to the flowfield measurements for expanded cooling hole geometries being unique to the literature, the testing facility used for these measurements was also unique in that both the external mainstream Mach number (Ma ∞ = 0.25) and internal coolant supply Mach number (Ma c = 0.3) were nearly matched. Results show that by expanding the exit of the cooling holes, the penetration of the cooling jet as well as the intense shear regions are significantly reduced relative to a round hole. Although the peak turbulence levels for all three hole geometries was nominally the same, the source of that turbulence was different. The peak turbulence level for both expanded holes was located at the exit of the cooling hole resulting from the expansion angle being too large. The peak turbulence level for the round hole was located downstream of the hole exit where the velocity gradients were very large.

Experimental and theoretical work concerning the application of ceramic components in small high temperature gas turbines has been performed for several years. The significance of some non-oxide ceramic materials for gas turbines in particular is based on their excellent high temperature properties. The application of ceramic materials allows an increase of the turbine inlet temperature resulting in higher efficiencies and a reduction of pollution emissions. The inherent brittleness of monolithic ceramic materials can be virtually reduced by reinforcement with ceramic fibers leading to a quasi-ductile behavior. Unfortunately, some problems arise due to oxidation of these composite materials in the presence of hot gas flow containing oxygen. At the Motoren- und Turbinen Union, München GmbH, comprehensive investigations including strength, oxidation, and thermal shock tests of several materials that seemed to be appropriate for combustor liner applications were undertaken. As a result, C/C, SiC/SiC, and two C/SiC-composites coated with SiC, as oxidation protection, were chosen for examination in a gas turbine combustion chamber. To prove the suitability of these materials under real engine conditions, the fiber reinforced flame tubes were installed in a small gas turbine operating under varying conditions. The loading of the flame tubes was characterized by wall temperature measurements. The materials showed different oxidation behavior when exposed to the hot gas flow. Inspection of the C/SiC-composites revealed debonding of the coatings. The C/C- and the SiC/SiC-materials withstood the tests with a maximum cumulated test duration of 90 hours without damage.

An experimental analysis of the slot film cooling performance downstream of a mixing jet has been performed. The mixing jet was blown into the main flow at a position five jet diameters downstream of the slot exit. Flow visualizations of the jet and of the cooling film are presented for blowing rates of 1 and 2 with jet momentum rates of 7 and 10. LDV-measurements in the symmetry plane show the strong perturbation acting on the cooling film downstream of the mixing jet injection. Heat transfer coefficients and adiabatic wall effectiveness were determined using electrical heater foils in combination with infrared-thermography. The results clearly show the decrease in film cooling performance in the region downstream of the mixing jet.

A unified expression for the spectrum of turbulence is developed by asymptotically matching known expressions for small and large wave numbers, and a formula for the one-dimensional spectral function which depends on the turbulence Reynolds number Re λ is provided. In addition, formulas relating all the length scales of turbulence are provided. These relations also depend on Reynolds number. The effects of free-stream turbulence on laminar heat transfer and pre-transitional flow in gas turbines are re-examined in light of these new expressions using our recent thoughts on an ‘effective’ frequency of turbulence and an ‘effective’ turbulence level. The results of this are that the frequency most effective for laminar heat transfer is about 1.3U/L e , where U is the free-stream velocity and L e is the length scale of the eddies containing the most turbulent energy, and the most effective frequency for producing pre-transitional boundary layer fluctuations is about 0.3U / η where η is Kolmogorov’s length scale. In addition, the role of turbulence Reynolds number on stagnation heat transfer and transition is discussed, and new expressions to account for its effect are provided.

A 3D Navier-Stokes code, together with the standard k-ϵ model with wall function approach, was used to investigate the flowfield in the vicinity of three different single scaled-up film-cooling holes. The hole geometries include a cylindrical hole, a hole with laterally expanded exit, and a hole with forward-laterally expanded exit. Comparisons of numerical results with detailed flowfield measurements of mean velocity and turbulent quantities are presented for a blowing ratio and density ratio of unity. Additionally, experimental data for different blowing ratios and a density ratio of about two are taken to perform validation of the code for adiabatic film-cooling effectiveness prediction. Results show that for both the round and the expanded hole geometries the code is able to capture all dominating flow structures of this jet in crossflow problem. However, discrepancies are found when comparing the flowfield inside the hole and at the hole exit. In particular, jet location at the hole exit differs significantly from measurement for the expanded hole geometries. For the adiabatic film-cooling effectiveness, it is shown that for round and expanded hole exits the intensity of the shear regions and the source of turbulence, respectively, have a strong influence on the predictive capability of the numerical code.

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system which provides a two-dimensional distribution of the film-cooling effectiveness in the nearfield of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the ejected coolant than the fanshaped hole which leads to higher laterally averaged film-cooling effectiveness. Coolant passage crossflow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

This paper presents the discharge coefficients of three film-cooling hole geometries tested over a wide range of flow conditions. The hole geometries include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered were the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the pressure ratio across the hole (up to 2). The results show that the discharge coefficient for all geometries tested strongly depends on the flow conditions (crossflows at hole inlet and exit, and pressure ratio). The discharge coefficient of both expanded holes was found to be higher than of the cylindrical hole, particularly at low pressure ratios and with a hole entrance side crossflow applied. The effect of the additional layback on the discharge coefficient is negligible.

A mathematical model to describe the coupled heat transfer in effusion-cooled combustor walls is presented. A heat balance with respect to the heat transfer with film cooling on the hot side, heat transfer on the cold side and heat exchange between the wall and the coolant inside the holes leads to a system of four equations for the unknown temperatures on the hot and cold side of the wall and for the inlet and exit temperature of the coolant. The model is validated with experimentally determined exit temperatures. A parameter study showing the influence of the major dimensionless parameters is performed.