Jens von Wolfersdorf

Time-Resolved Heat Transfer Measurements on the Tip Wall of a Ribbed Channel Using a Novel Heat Flux Sensor—Part I: Sensor and Benchmarks

A novel heat flux sensor was tested that allows for time-resolved heat flux measurements in internal ribbed channels related to the study of passages in gas turbine blades. The working principle of the atomic layer thermopile (ALTP) sensor is based on a thermoelectric field created by a temperature gradient over an yttrium-barium-copper-oxide (YBCO) crystal (the transverse Seebeck effect). The sensors very fast frequency response allows for highly time-resolved heat flux measurements up to the 1 MHz range. This paper explains the design and working principle of the sensor, as well as the benchmark-ing of the sensor for several flow conditions. For internal cooling passages, this novel sensor allows for highly accurate, time-resolved measurements of heat transfer coefficients, leading to a greater understanding of the influence of fluctuations in temperature fields.

Experimental Investigations on Transpiration Cooling for Scramjet Applications Using Different Coolants

The extremely high heat loads within a scramjet combustor require the use of high-temperature materials combined with efficient cooling concepts. A promising technique is the application of transpiration cooling to ceramic matrix composite materials. A supersonic hot-gas-flow test facility is used to investigate this cooling method. The carbon/carbon samples tested have porosities of about e = 11%. The airflow is electrically heated up to 1120 K total temperature with a total pressure of ≈3 bar and is accelerated to a Mach number of 2.1 within the test channel. Air, argon, and helium are used as coolants for blowing ratios from 0 to 1 %. The surface temperature of the porous wall is measured via thermocouples and infrared thermography. Pressure and mass-flow measurements are used to analyze the throughflow characteristics of the porous materials at various temperature conditions. An approach based upon simplified analytical models is presented to analyze the experimental data of throughflow behavior and cooling efficiency. The simplified thermal model is used to analyze the effect of fluid property variations with temperature on pressure loss for different coolants and shows good agreement with the experimental data.

Experimental Study on Combustion Mode Transition in a Scramjet with Parallel Injection

Mode transition from weak combustion to strong combustion or vice versa in a scramjet engine is a critical phenomenon in designing such engines, because the thrust of each mode varies considerably. The mode transition is supposed to interact strongly with a so-called pseudo-shock wave or shock train. In order to control vehicles with scramjet engines, it is, therefore, essential to understand mode transition. Several studies concerning this phenomenon have been conducted and most of them used the wall fuel injection method, where the disturbance of the boundary layer due to the wall injection could not be avoided. In order to prevent this disturbance, a parallel injection method was used in this study. The experiments were conducted using the supersonic combustion facilities of

The Effect of Turning Vanes on Pressure Loss and Heat Transfer of a Ribbed Rectangular Two-Pass Internal Cooling Channel

Gas turbine blades are usually cooled by using ribbed serpentine internal cooling passages, which are fed by extracted compressor air. The individual straight ducts are connected by sharp 180 deg bends. The integration of turning vanes in the bend region lets one expect a significant reduction in pressure loss while keeping the heat transfer levels high. Therefore, the objective of the present study was to investigate the influence of different turning vane configurations on pressure loss and local heat transfer distribution. The investigations were conducted in a rectangular two-pass channel connected by a 180 deg sharp turn with a channel height-to-width ratio of HIW=2. The channel was equipped with 45 deg skewed ribs in a parallel arrangement with e/d h = 0.1 and P/e = 10. The tip-to-web distance was kept constant at W el /W = 1. Spatially resolved heat transfer distributions were obtained using the transient thermochromic liquid crystal technique. Furthermore static pressure measurements were conducted in order to determine the influence of turning vane configurations on pressure loss. Additionally, the configurations were investigated numerically by solving the Reynolds-averaged Navier― Stokes equations using the finite-volume solver FLUENT. The numerical grids were generated by the hybrid grid generator CENTAUR. Three different turbulence models were considered: the realizable k-e model with two-layer wall treatment, the k-ω-SST model, and the v 2 -f turbulence model. The results showed a significant influence of the turning vane configuration on pressure loss and heat transfer in the bend region and the outlet pass. While using an appropriate turning vane configuration, pressure loss was reduced by about 25%, keeping the heat transfer at nearly the same level in the bend region. An inappropriate configuration led to an increase in pressure loss while the heat transfer was reduced in the bend region and outlet pass.

Shape Optimization of Cooling Channels Using Genetic Algorithms

A shape optimization method for convective cooling channels within a two-dimensional heat conduction region is presented. This method combines genetic algorithms with a point heat sink approach that is used to model the heat removal of the cooling channels during the optimization process. The method can be easily combined with the Finite Element Method (FEM) for the calculation of the optimized temperature field distribution.

Experimental Investigations of Scramjet Combustor Characteristics

The design of scramjet engines requires systematic experimental data. Hence, in the present study, a scramjet combustor model was investigated experimentally using the facil- ities at the Institute of Aerospace Thermodynamics, UniversitStuttgart, Germany. The facilities can deliver a continuous supersonic flow of Mach number 2.1 with a maximum total temperature of 1500K and a maximum total pressure of 1MPa. This corresponds to a flight Mach number of 5 at an altitude of about 30km. In the present experiments, hydrogen was injected using a central injector concept that enhances mixing efficiency by inducing streamwise vortices. Parameters used in this study include injector type, injection position, combustor opening angle, total temperature and equivalence ratio. Measurements of static wall pressure distributions and high speed pictures of the flow field indicated the presence of two distinctive combustion modes. These were dependent on running condi- tions. Weak combustion, which is accompanied by small pressure rise, showed a detached flame. Strong combustion, where the flame is attached to the injector, showed a higher pressure rise. The transition between the two modes was very sensitive to the test con- ditions at the investigated Mach number. The tansition mechanism was supposed to be a detonation wave travelling upstream to the trailing edge of the injector. After mode transition occured, a stable combustion was achieved, where the central injector acted as a kind of flameholder. Due to the geometric blockage of the injector and the growing bound- ary layer the static pressure in the isolator increased. In all test cases this pressure rise causes a pre-combustion shock train structure. In the present study the central injection concept was found to have good mixing and combustion capabilities. Especially if strong combustion is active and combustor height is sufficient to reduce geometric blockage effects.

Thermocouple Thermal Inertia Effects on Impingement Heat Transfer Experiments Using the Transient Liquid Crystal Technique

The transient liquid crystal technique is widely used for impingement heat transfer experiments. Additionally, due to the difficulty of producing pure temperature steps in the flow, many authors assumed the fluid temperature evolution as a series of step changes using Duhamels superposition theorem. However, for small impingement configurations where the jets are fed from the same plenum chamber, and hence flow velocities are relatively small, thermal inertia of commercial thermocouples causes a delay, lagging from the real plenum temperature history. This paper investigates thermal inertia characteristics of thermocouples and their effect on the calculation of impingement heat transfer coefficient. Several thermocouples with exposed junction and different wire diameter were considered over a range of plenum flow conditions typically found in impingement heat transfer experiments. The effect of thermocouple time constant on the evaluation of the heat transfer rate was investigated in a narrow channel consisting of five inline impingement jets. The results indicated a significant effect of thermocouple response on the stagnation point region heat transfer, while lower local heat transfer rates are negligibly affected as liquid crystal signals appear later in time and the driving gas temperature history has a smaller influence on the evaluated data.

Numerical Study of Supersonic Combustion Processes with Central Strut Injection

The combustion chamber is shown in Fig. 1. It has a rectangular cross section with width w 40 mm and constant isolator height h 35:4 mm. Downstream distance, x, is measured from the Laval nozzle throat where x 0. Air enters the chamber from the left at a Mach number of M 2:1 at isolator entry. The combustor and diverging sections open at angles 1 1 and 2 2 , respectively. The outflow static pressure is p1 0:96 bar. The facility operates continuously at stagnation conditions T0;air 1400 K and p0;air 4 bar. For brevity, in the following, wall pressures in the symmetry plane are only shown for the upper wall. Our injector is shown in Fig. 1b. Chun et al. [1] provide further details about the facility and Gerlinger et al. [2,3] provide further details on the injector. Because of exploitation of symmetry, only half the channel width is modeled. H2 is injected through the horizontal slots (see Fig. 1b). At design conditions, the injection Mach number is MH2 2:6. However, in practice, this is probably not achieved, due to geometric deformation. As a conservative estimate we chose MH2 1:0. The validity of this choice is checked in Sec. V.E. Four combustion modes, i.e., blowoff, weak combustion, strong combustion, and thermal choking were observed. Because of flow overexpansion, all modes feature a shock train in proximity of the channel outflow. Another shock train develops in the isolator, due to injector displacement effects. In flight, strong combustion is desired. Here, the injector acts as the flame holder and heat release is high. For weak combustion the fuel–air mixture is ignited in the outflow shock train. The flame is detached from the injector and heat release is significantly lower. If, in these conditions, the exit pressure is lowered sufficiently, the outflow shock train disappears, and combustion ceases (blowoff). For excessive fuel injection, the flow becomes thermally choked. The combustion mode is determined by the geometry, stagnation conditions, ambient pressure, and the global equivalence ratio (see also Scheuermann et al. [4]).

Heat Transfer Technology for Internal Passages of Air-Cooled Blades for Heavy-Duty Gas Turbines

Abstract: The present review paper, although far from being complete, aims to give an overview about the present state of the art in the field of heat transfer technology for internal cooling of gas turbine blades. After showing some typical modern cooled blades, the different methods to enhance heat transfer in the internal passages of air‐cooled blades are discussed. The complicated flows occuring in bends are described in detail, because of their increasing importance for modern cooling designs. A short review about testing of cooling design elements is given, showing the interaction of the different cooling features as well. The special focus of the present review has been put on the cooling of blades for heavy‐duty gas turbines, which show several differences compared to aero‐engine blades.

An Experimental and Numerical Study of Heat Transfer and Pressure Loss in a Rectangular Channel With V-Shaped Ribs

An experimental and numerical study was conducted to determine the thermal performance of V-shaped ribs in a rectangular channel with an aspect ratio of 2:1. Local heat transfer coefficients were measured using the steady state thermochromic liquid crystal technique. Periodic pressure losses were obtained with pressure taps along the smooth channel sidewall. Reynolds numbers from 95,000 to 500,000 were investigated with V-shaped ribs located on one side or on both sides of the test channel. The rib height-to-hydraulic diameter ratios (e/D h ) were 0.0625 and 0.02, and the rib pitch-to-height ratio (P/e) was 10. In addition, all test cases were investigated numerically. The commercial software FLUENT™ was used with a two-layer k-s turbulence model. Numerically and experimentally obtained data were compared. It was determined that the heat transfer enhancement based on the heat transfer of a smooth wall levels off for Reynolds numbers over 200,000. The introduction of a second ribbed sidewall slightly increased the heat transfer enhancement whereas the pressure penalty was approximately doubled. Diminishing the rib height at high Reynolds numbers had the disadvantage of a slightly decreased heat transfer enhancement, but benefits in a significantly reduced pressure loss. At high Reynolds numbers small-scale ribs in a one-sided ribbed channel were shown to have the best thermal performance.