Impingement cooling in airfoils cooling cavities, solely or combined with film and convective cooling, is a common practice in gas turbines. Depending on the cooling cavity design, the mass flow rate through individual crossover holes could vary significantly in the flow direction thus creating jets of different strengths in the target cavity. This jet flow variation, in turn, creates an impingement heat transfer coefficient variation along the channel. A test section, simulating two adjacent cooling cavities on the trailing side of an airfoil, is made up of two channels with trapezoidal cross-sectional areas. On the partition wall between the two channels, eleven crossover holes create the jets. Two distinct exit flow arrangements are investigated — a) jets, after interaction with the target surface, are turned towards the target channel exit axially and b) jets are exited from a row of racetrack-shaped slots along the target channel. Flow measurements are reported for individual holes and heat transfer coefficients on the eleven target walls downstream the jets are measured using the steady-state liquid crystal thermography technique. Smooth as well as rib-roughened target surfaces with four rib geometries (0°,45°, 90° and 135° rib angles) are tested. Correlations are developed for mass flow rate through each crossover hole for cases with different number of crossover holes, based on the pressure drop across the holes. Heat transfer coefficient variations along the target channel for all rib geometries and flow conditions are reported for a range of 5000 to 50000 local jet Reynolds numbers. Major conclusions of this study are: 1) A correlation is developed to successfully predict the mass flow rates through individual crossover holes for geometries with six to eleven crossover holes, based on the pressure drop across the holes, 2) impingement heat transfer coefficient correlates well with the local jet Reynolds number for both exit flow arrangements, and 3) the case of axial flow in the target channel exiting from the channel end, at higher jet Reynolds numbers, produced higher heat transfer coefficients than those in the case of flow exiting through a row of slots along the target channel opposite to the crossover holes.

Accurate temperature prediction of turbine blades for gas turbine is very important to assure the life-span of the blade under a hostile hot gas environment and intense centrifugal force. Therefore, there have been a number of studies carried out to clarify the cooling performance of serpentine cooling channel inside a turbine blade for gas turbine, however, it remains to be quite difficult to make an accurate numerical prediction of the performance. Apart from the effects of disk rotation as well as large temperature gradient near the wall, such a poor predictability can be attributed to the complicated vortical motions caused by the rib-roughened cooling channel whose cross-sectional shape varies along the channel and by the existence of u-bends. Furthermore, since the cooling channel inside a real turbine blade usually has a curved or S-shaped inlet, which may induce flow separation as well as swirl developed in the inlet, it can be imagined that the flow and heat transfer inside the cooling channel is likely to become much more complicated than that with a straight inlet. Despite this situation, only few studies are made in order to examine the flow and heat transfer characteristics inside the cooling channel with s-shaped inlet. Accordingly, this study aims at detailed experimental and numerical investigations on the flow and heat transfer characteristics of a realistic serpentine rib-roughened cooling channel with an s-shaped inlet, which is modeled from an actual HP turbine blade for gas turbine. This study employs a transient TLC (Thermochromic Liquid Crystal) technique to measure the heat transfer characteristics, along with the flow visualization on the inner surface of the channel using oil mixed with titanium powder. Note that a special focus in this flow visualization is placed on the area of s-shaped inlet. As for the flow measurement, 2D-PIV (Particle Image Velocimetry) method is used to understand time-dependent vortical structures of the flow field that can have significant impacts on the heat transfer. RANS-based numerical simulation is also executed to predict the heat transfer distribution on the inner surface of the cooling channel.

This paper describes testing an axial turbine response when driven by a Rotating Detonation Combustor (RDC). A T63 (C20-250) gas turbine is modified by replacing the combustor with a RDC. The stator vanes of the T63 are heavily instrumented for measurement of flow enthalpy and pressure. The engine is run at multiple power levels with the stock combustor using JetA and hydrogen fuel. The engine is then modified to have an open loop configuration, and is run with both the RDC and the stock combustor hardware with hydrogen fuel. Temperature pattern factor, flow unsteadiness, and turbine component efficiency are measured for all setups. High speed pressure transducers show substantially higher unsteadiness generated by the RDC than the conventional combustor. RDC turbine component efficiencies are compared to the conventional combustor. Results suggest that RDC unsteadiness does not significantly impact turbine efficiency.

A robust 1D film hydrodynamic model has been sequentially coupled with a 1D core gas model and used to predict the instantaneous mean core gas speed, film interface shear stress and liquid film distribution within an idealised bearing chamber. This novel approach to aero-engine bearing chamber simulation provides a predictive tool that can be used for the fast and reliable exploration of a set of bearing chamber design and operating conditions characterised by the: chamber dimensions, air/oil fluid properties, shaft speed, sealing air flows, oil feed rates and sump scavenge ratios. A preliminary validation of the model against available bearing chamber flow measurements from literature shows good agreement. The model represents a significant step change in predictive capabilities for aero-engine oil system flows compared to previous semi-empirical models. The bearing chamber is idealised as a one-dimensional (2D) domain with a predominantly azimuthal flow in both the rotational oil film and core gas such that axial components may be ignored. A 1D system of depth-averaged film hydrodynamics equations is used to predict oil film thickness and mean speed distributions in the azimuthal direction under the influence of interface shear, gravity, pressure gradient and surface tension forces. The driving shear stress in the film model is obtained from the 1D core-gas model based on an azimuthal gas momentum conservation equation which is coupled to the film model through the interface shear stress and film interface velocity.

Printed circuit heat exchangers (PCHEs) have an important role in supercritical CO 2 (sCO 2 ) Brayton cycles because of their small footprint and the high level of recuperation required for this power cycle. Compact heat exchangers like PCHEs are a rapidly evolving technology, with many companies developing various designs. One technical unknown that is common to all compact heat exchangers is the flow distribution inside the headers that affects channel flow uniformity. For compact heat exchangers, the core frontal area is often large compared with the inlet pipe area, increasing the possibility of flow maldistribution. With the large area difference, there is potential for higher flow near the center and lower flow around the edges of the core. Flow maldistribution increases pressure drop and decreases effectiveness. In some header geometries, flow separation inside the header adds to the pressure drop without increasing heat transfer. This is the first known experiment to test for flow maldistribution by direct velocity measurements in the headers. A PCHE visualization prototype was constructed out of transparent acrylic for optical flow measurements with Particle Image Velocimetry (PIV). The channels were machined out of sheets to form many semi-circular cross sections typical of chemically-etched plates used in PCHE fabrication. These plates were stacked and bolted together to resemble the core geometry. Two header geometries were tested, round and square, both with a normally-oriented jet. PIV allows for velocities to be measured in an entire plane instantly without disturbing the flow. Small particles of approximately 10 micrometers in diameter were added to unheated water. The particles were illuminated by two laser flashes that were carefully timed, and two images were acquired with a specialized digital camera. The movement of particle groups was detected by a cross-correlation algorithm with a result of about 50k velocity measurements in a plane. The velocity distribution inside the header volume was mapped using this method over many planes by traversing the PCHE relative to the optical equipment. The level of flow maldistribution was measured by the spatially-changing velocity coming out of the channels. This effect was quantified by the coefficient of variation proposed by Baek et al. The relative levels of flow maldistribution in the different header geometries in this study were assessed. With highly-resolved velocity measurements, improvements to header geometry to reduce flow maldistribution can be developed.

Numerical simulation results of a single jet premixed combustion system at atmospheric pressure are compared against comprehensive particle image velocimetry (PIV) flow measurements and Raman scattering temperature measurements for natural gas and hydrogen fuels. The simulations were performed on hexahedral meshes with 1–5 million elements. RANS calculations were carried with the k-ε realizable turbulence model. Combustion was modeled using the Flamelet Generated Manifold model (FGM) and detailed chemistry. Both the flame position and flame liftoff predicted by the FGM were in reasonable agreement with experiments for both fuels and showed little sensitivity to heat transfer or radiation modeling. The detailed chemistry calculation predicts the temperature gradients along the jet centerline accurately and compares very closely with the Raman scattering measurements. The much closer agreement of the jet axial velocity and temperature profiles with experimental values, coupled with the significantly protracted presence of intermediates in the detailed chemistry predictions, indicates that the impact of nonequilibrium intermediates on very lean natural gas flames is significant.

A rotating test rig, reproducing a rotor-stator cavity with an axial admission system, has been exploited for an experimental investigation on the internal flow field and its effect on heat transfer on the stator side. Working conditions were varied in a wide range of rotating velocities and superposed mass flow rates. 2D PIV flow measurements were performed in order to obtain a radial distribution of the tangential velocity, results were used to validate numerical simulations aimed at understanding the admission system effect on the swirl distribution. Heat transfer coefficient distribution along the stator disk has been evaluated performing a steady state technique exploiting Thermo-chromic Liquid Crystals (TLC). Tests have been performed varying the superposed mass flow rate up to reaching the condition of cavity completely sealed, further increase of the mass flow rate showed to reduce the effect of the rotation. Working conditions were set in order to investigate cases missing in open literature, however few tests performed in similarity with other researches provided comparable results.

Currently fast response aerodynamic probes are widely used for advanced experimental investigations in turbomachinery applications. The most common configuration is a virtual three hole probe. This solution is a good compromise between probe dimension and accuracy. Several authors have attempted to extend the capabilities of these probes in terms of bandwidth and operating conditions. Even though differences exist between the solutions in the literature, all of the designs involve the positioning of a dynamic pressure sensor close to the measurement point. In general terms, the higher the frequency response, the more the sensor is exposed to the flow. This physical constraint puts a limit on the probe applicability since the measurement conditions have to comply with the maximum allowed operating conditions of the sensor. In other applications, when the conditions are particularly harsh and a direct measurement is not possible, a waveguide probe is commonly used to estimate the local pressure. In this device the sensor is connected to the measurement point through a transmitting duct which guarantees that the sensor is operating in a less critical condition. Generally, the measurement is performed through a pressure tap and particular attention must be paid to the probe design in order to have an acceptable frequency response function. In this study, the authors conceived, developed and tested a probe which combines the concept of a fast response aerodynamic pressure probe with that of a waveguide probe. Such a device exploits the benefits of having the sensor far from the harsh conditions while maintaining the capability to perform an accurate flow measurement.

This paper presents an analytical approach to quantify the uncertainty of velocity-, angle- and turbulence intensity measurements obtained by hot-wire anemometry. Based on the given results of the model, the paper points out the most critical parameters influencing the quality of the flow measurement. The proposed approach is divided into three subsections, each accounting for the uncertainties due to the measurement equipment, probe calibration and measurements in the test-rig respectively. For every subsection the propagation of uncertainties is stated in the form of analytical formulas starting with the measurands, e.g. voltage, pressure, temperature, and ending with velocity, angle and turbulence intensity. The approach is applied to data, obtained by a triple hot-wire probe, on two-dimensional traverses in a 1.5-stage cold air turbine. The calibration process and, in particular, the temperature and pressure deviation between calibration and measurement are identified as having a major impact on the measurement uncertainty. With respect to the Aachen test-rig the uncertainty in velocity measurements is lower than 3.5%. The investigation of the turbulence intensity uncertainty yields an increase in uncertainty with increasing turbulence intensity but an insensitivity towards pitch- and yaw-angle uncertainty.

The full-field endwall Nusselt number ( Nu ) distributions and flow field are presented respectively using steady-state infrared thermography and particle image velocimetry (PIV) for the two-pass parallelogram channels with attached and detached transverse ribs. These square transverse ribs on two opposite channel endwalls are in-line arranged with rib-height to duct-height ratio of 0.1 and rib-pitch to rib-height ratio of 10. For the detached ribs, the detached distance between rib and channel endwall is 0.38 rib height. With the measurements of Fanning friction factor ( f ), heat transfer distributions and flow field features, the thermal performance factors (TPF) for the attached and detached rib cases are comparatively examined. A set of Nu , f and TPF results with the associated flow measurements at the test conditions of 5,000≤ Re ≤20,000 is selected to disclose the differential heat transfer enhancement mechanisms and heat transfer efficiencies between the attached and detached ribbed channels. Empirical correlations evaluating the endwall area-averaged Nusselt numbers ( Nu ) and f factors are devised to assist the relevant design activities.

In this study, we experimentally investigate both the intrinsic instability characteristics and forced response to transverse acoustic excitation of a non-reacting, swirling flow for application to combustion instability in annular gas turbine engines. The non-axisymmetry of the velocity field is quantified using an azimuthal mode decomposition of the time-averaged velocity field that shows that (1) the flow field is largely axisymmetric, (2) axisymmetry decreases with downstream distance, and (3) forcing does not significantly alter the time-averaged shape of the flow field. The flow field is analyzed in a companion linear stability analysis that shows that the most unstable modes in the flow field are m =−1 and m =−2, which agrees with the experimental observations and shows that the intrinsic dynamics of this flow field are non-axisymmetric with respect to the jet axis. The linear stability analysis captures the spatial variation of mode strength for certain modes, particularly mode m =−1, but there are some deviations from the experimental results. Most notably, these deviations occur for mode m =0 at radii away from the jet axis. Experimental results of the forced response of the flow indicate that the intrinsic instability characteristics of the flow field have an impact on the forced-response dynamics. Response of the flow field to a velocity anti-node in a standing transverse acoustic field shows non-axisymmetric vortex rollup and the dominance of the m =−1 and m =1 azimuthal modes in the fluctuating flow field. In the presence of a pressure anti-node, the m =0 mode of the fluctuating flow field is very strong at the jet exit, indicating an axisymmetric response, and ring vortex shedding is apparent in the flow measurements from high-speed PIV. However, further downstream, the strength of the axisymmetric mode decreases and the m =−1 and m =1 modes dominate, resulting in a tilting of the vortex ring as it convects downstream. Implications for flame response to transverse acoustic fields are discussed.

Time-accurate numerical simulations were conducted on the aft-loaded L1A low pressure turbine airfoil at a Reynolds number of 22,000 (based on inlet velocity magnitude and axial chord length). This flow condition produces a non-reattaching laminar separation zone on the airfoil suction surface. The numerical code TURBO is used to simulate this flow field as an Implicit Large Eddy Simulation. Generally good agreement was found when compared to experimental time-averaged and instantaneous flow measurements. The numerical separation zone is slightly larger than that in the experiments, though integrated wake loss values improved from RANS-based simulations. Instantaneous snapshots of the numerical flow field showed the Kelvin Helmholtz instability forming in the separated shear layer and a large-scale vortex shedding pattern at the airfoil trailing edge. These features were observed in the experiments with similar sizes and vorticity levels. Power spectral density analyses revealed a global passage oscillation in the numerics that was not observed experimentally. This oscillation was most likely a primary resonant frequency of the numerical domain.

Large-scale (that is over the kilometre scale of a whole wind farm) and small-scale (over the metre scale of an individual wind turbine) wind flow measurements are made with a mobile-based LIDAR system. The large-scale measurements detail the multiple wakes in the wind farm, including both single and double wakes, that result in up to 70% less power generation for a given wind direction, and 2.5% loss of the wind farm’s annual energy yield. The small-scale measurements show that there is a non-uniform work extraction of the turbine across the vertical extent of the wind turbine rotor. This non-uniform work extraction is accompanied by an upward pitching of the flow that is as much as 50° immediately downstream of the turbine and reduces to 10° two diameters downstream. Measurements with the mobile-based LIDAR system are made in both complex and flat terrains. A comparison of the wake profiles show that whereas in complex terrain the profiles are self-similar up to two-and-half rotor diameters downstream, this is not the case in flat terrain. It is shown that these measurements, which are made at the full-scale Reynolds number in the field, may be useful to support the development of wake flow prediction tools.

This paper presents the results of a wide experimental study on an H-type Vertical Axis Wind Turbine (VAWT) carried out at Politecnico di Milano in the frame of the national founded project PRIN 2009. The experiments were carried out in the large-scale wind tunnel of the Politecnico di Milano, where real-scale wind turbines for micro generation can be tested in full similarity conditions. Integral torque and thrust measurements were performed, as well as detailed aerodynamic measurements to characterize the flow field generated by the turbine downstream of the rotor. The machine was tested in both confined and non-confined environment, to highlight the effect of wind tunnel blockage on the aerodynamics and performance of the VAWT under investigation. The experimental results, compared with the blockage correlations presently available, suggest that specific correction models should be developed for VAWTs. The experimental thrust and power curves of the turbine, derived from integral measurements, exhibit the expected trends with a peak power coefficient of about 0.28 at tip-speed ratio λ = 2.5. Detailed flow measurements, performed in three conditions for λ equal to 1.5, 2.5 and 3.5, show the fully three-dimensional character of the wake, especially in the tip region where a non-symmetrical wake and tip vortex are found. The unsteady evolution of the velocity and turbulence fields further highlight the effect of aerodynamic loading on the wake unsteadiness, showing the time-dependent nature of the tip vortex and the onset of a non-symmetric dynamic stall for λ lower than 2.

Flow measurements taken on the rotors of turbomachines are of great value for improved understanding and advancement of design techniques. On-rotor experiments have been limited in the past because of instrumentation problems, especially with the data transmission system. Recent advances in miniature electronic systems and transducer technology have produced a renewed interest in this area. These considerations are discussed, and research on a telemetry-type data transmission system is described with experimental verification using a strain gage on an axial-flow fan blade.

Similar to the Doppler technique the principle of this method is also based on light scattered by small particles as they are normally contained in every fluid. Two light beams (laser beams are not necessary) are focused in two very small light spots in the measuring volume. Particles passing both the spots emit two light pulses which can be detected by a photomultiplier. From the time interval between these two pulses and from the known spots distance the flow velocity can be calculated. A statistical analysis of many signals, taken at the same treasuring point, allows calculation of the mean velocity, the mean flow angle and the turbulence degree of the velocity components in a plane normal to the beam axis. Due to the light intensity of the light spots (10 2 to 10 3 times higher than the intensity in a Doppler probe volume under similar conditions) a sufficient signal-to-noise ratio is achieved so that velocity measurements even in unseeded air flows are possible up to 500 m/s in backscattering by using a 5 mW laser. Tests were carried out in a supersonic wind tunnel up to a Mach number of 2.2. For measurements in the rotor blade channels of turbomachines a specific trigger optic has been installed.

Freezing behavior, pumpability, and temperature profiles for aviation turbine fuels were measured in a 190-liter tank, to simulate internal temperature gradients encountered in commercial airplane wing tanks. Two low-temperature situations were observed. Where the bulk of the fuel is above the specification freezing point, pumpout of the fuel removes all fuel except a layer adhering to the bottom chilled surfaces, and the unpumpable fraction depends on the fuel temperature near these surfaces. Where the bulk of the fuel is at or below the freezing point, pumpout ceases when solids block the pump inlet, and the unpumpable fraction depends on the overall average temperature.

Time-mean flow measurements of turbulent boundary layer development on the convex surface of an outlet stator blade in a single-stage axial compressor are presented. There is no evidence of logarithmic wall similarity at blade chord Reynolds numbers from 3 × 10 4 to 2 × 10 5 , and its absence appears due to the combined effects of low Reynolds number, large positive pressure gradient and rapidly changing boundary conditions. Conventional skin friction laws compare very poorly with experiment. The performance of local equilibrium and entrainment-type calculation methods is examined and serious errors are found to develop at blade Reynolds numbers below 10 5 . The best results are obtained from a lag-entrainment method of Green, Weeks and Brooman, which can be recommended for predicting axial turbomachine blade boundary layers at moderate Reynolds number.

Detailed examination of flow measurements over concave pressure surfaces suggests that interaction of Taylor-Goertler vorticity with mainstream turbulent exerts only limited influence in enhancing laminar boundary-layer heat transfer. While transition is primarily controlled by the Launder laminarisation criterion, the Goertler number at which it subsequently occurs is not solely determined by turbulence intensity. Adoption of K >2.5.10 ± as a design criterion for the pressure surfaces of turbine blades would appear to have significant advantages in terms of reduced heat transfer, increased lift, and lower aerodynamic drag.

Detailed flow measurements made in a 1-m dia shrouded centrifugal impeller running at 500 rpm are presented. All 3 mutually perpendicular components of relative velocity and rotary stagnation pressures were measured on 5 cross-sectional planes between the inlet and the outlet, using probes which were traversed within the rotating impeller passage. The reduced static pressures were also calculated from these flow measurements. The measurements were made for an impeller flow rate corresponding to approximately zero incidence at the blade leading edges. Shroud boundary layer separation and secondary flow were observed to lead to the formation of a wake in the suction-side/shroud corner region. It is concluded that the turbulent mixing associated with the shroud boundary layer separation and the strength of the secondary flow strongly influence the size and location of the wake respectively.