The transient Thermochromic Liquid Crystal (TLC) method is applied to determine the distribution of the local heat transfer coefficients using a configuration with parallel cooling channels at an engine relevant Reynolds number. The rectangular channels with a moderate aspect ratio and a high length-to-diameter ratio are equipped with one-sided oblique ribs with high blockage, which is a promising configuration for turbine near wall cooling applications. In this arrangement, the three inner channels should experience same flow and thermal conditions. Numerical simulations are performed to substantiate this assumption. The symmetric single channels are sprayed with narrowband TLC with various indication temperatures. Multiple experiments were conducted. All start at ambient conditions before the fluid is heated up to several temperatures between 46°C and 73°C. The results show that the determined local heat transfer coefficients and therefore the Nusselt numbers vary significantly for the different experimental conditions especially at locations of high heat transfer coefficient behind the ribs. A simplified procedure with respect to measurement uncertainties is applied to enable an easy and fast valuation on the data quality. This might be used within the data reduction analysis for such experiments directly. The approach is illustrated using the obtained experimental data.

Patching is a high-tech repair procedure that is very adequate for compressor blisks with larger damages. This repair concept has the advantage that the added patch provides the same mechanical strength as the parent material of the blade and the initial aerodynamic contour of the blade is fully restored. However, the welding process locally induces stresses in the heat affected zone at the patch-to-blisk interface. These welding residual stresses influence the fatigue life of the repaired blade and have to be considered during the design phase of patch repairs. In this work, we contribute to the design of patch repairs by introducing a numerical simulation to predict weld-induced stresses in repaired compressor blades. Therefore, a finite element model is developed that includes sequential thermal and mechanical analyses of blisk blades. The temperature field caused by the welding torch is determined by performing a transient heat transfer analysis. The model also reflects the changes in the geometry due to the additional patch material and subsequent re-contoured patch. Different patch geometries are evaluated and compared in terms of their resulting stress levels. Basically, two kinds of patch geometries with long and short welding seams are studied. The stationary stress distribution of the repaired blade results from the superposition of residual stresses with steady stresses due to rotational and pressure forces. Thus, we provide the basis for a new fatigue assessment of the repaired blade considering the residual stress level in the patch-to-blisk interface.

A realistic internal cooling system of a turbine blade includes both ribs and pin-fins inside the passages to enhance the heat transfer. However, the majority studies in the open literature assessing the heat transfer characteristics on a simplified cooling model by examining ribbed-roughen passages and pin-finned passage separately. This work presents the high-resolution heat transfer coefficients of a scaled realistic turbine blade internal cooling design. The cooling system, using a 3D-printed plastic material, consists of an S-shaped inlet, four serpentine passages (three U-bends) of variable aspect ratio, and the trailing edge ejection. Angled ribs are implemented inside the passages and the elongated fins and pins are used near the trailing edge. Two dust holes are realized on the blade tip, the injections are individually controlled to reflect the realistic coolant flow rate variation inside the entire internal cooling system. The tests are conducted at two Reynolds number, 45,000 and 60,000 based on the hydraulic diameter of the inlet passage. Transient heat transfer technique using thermochromic liquid crystal is applied to obtain the detailed heat transfer characteristic inside the cooling channel. The local and averaged Nusselt numbers are also compared with the correlations in the open literature. This paper provides gas turbine designers the difference of local heat transfer distributions between the realistic and simplified internal cooling designs.

A smooth two-pass internal gas turbine cooling channel is numerically investigated. Transient conjugated non-rotating and rotating URANS simulations are executed. The transient thermochromic liquid crystal (TLC) approach is supported with these simulations as the temporally changing rotational buoyancy effects can be examined. The Reynolds number is 25,000, the rotation number is 0.24 and the initial buoyancy number is 0.63 (according to an inlet-to-wall density ratio of 0.23). As heat is transferred, the temperatures and heat fluxes change with increasing time, and so do the local buoyancy effects. The computational results are evaluated as averaged segmental values. They are compared to the experimental results from literature that have been determined for constant wall temperature experiments (various experimental runs with different constant wall temperatures). Especially in the first passage, there is a good agreement between the numerically gained results and the experimental data. The more complex flow inside the bend leads to more diverse characteristics and the second passage is only slightly effected.

The present study deals with the application of the transient thermochromic liquid crystal (TLC) technique in a flow network of intersecting circular passages as a potential internal turbine component cooling geometry. The investigated network consists of six circular passages with a diameter d = 20mm that intersect coplanar at an angle θ = 40°, the innermost in three, the outermost in one intersection level. Two additional non-intersecting passages serve as references. Such a flow network entails specific characteristics associated with the transient TLC method that have to be accounted for in the evaluation process: the strongly curved surfaces, the mixing and mass flow redistribution at each intersection point, and the resulting gradients between the wall and passage centerline temperatures. All this impedes the choice of a representative fluid reference temperature, which results in deviations using established evaluation methods. An alternative evaluation approach is introduced, which is supported by computational results obtained from steady-state three-dimensional RANS simulations using the SST turbulence model. The presented analysis uncouples local heat transfer coefficients from actually measured local temperatures but uses the time information of the thermocouples instead that represents the fluid temperature step change and evolution along the passages. This experimental time information is transferred to the steady-state numerical bulk temperatures, which are finally used as local references to evaluate the transient TLC experiments. As effective local mass flow rates in the passage sections are considered, the approach eventually allows for a conclusion whether heat transfer is locally enhanced due to higher mass flow rates or the intersection effects.

A series of experiments were conducted to investigate the detailed heat transfer characteristics of a large scaled model of a turbine blade internal cooling system. The cooling system has one passage in the leading edge and a triple passage for the remained region with two U-bends. A large scaled model (2 times) is designed to acquire high resolution measurement. The similarity of the test model was conducted with Reynolds number at the inlet of the internal cooling system. The model is designed to simulate the flow at engine condition including film extractions to match the changes in flowrates through the internal cooling system. Also, 45 deg ribs were installed for heat transfer enhancement. The experiments were performed varying Reynolds number in the range of 20,000 to 100,000 with and without ribs under stationary condition. This study employs transient heat transfer technique using thermochromic liquid crystal (TLC) to obtain full surface heat transfer distributions. The results show the detailed heat transfer distributions and pressure loss. The characteristics of pressure loss is largely dependent on the changes in cross-sectional area along the passages, the presence of U-bends and the extraction of coolant flow through film holes. The local and area averaged Nusselt number were compared to available correlations. Finally, the thermal performance counting the heat transfer enhancement as well as pressure penalty is presented.

The present study focuses on evaluating the thermal performance of the realistic leading edge cooling passage of a turbine blade, experimentally. The detailed heat transfer distributions and pressure loss of the cooling passage are measured with and without rib turbulators for direct comparison. The experimental model is representing the realistic leading edge cooling passage, which has a converging duct with triangular cross sectional shape. Therefore, the aspect ratio and the cross sectional area of the triangular passage vary significantly along the coolant flow direction. A Perspex test model is designed at 1.3 times large scale to meet Reynolds number similarity. The test model is installed with 45-degree rib turbulators at constant space-to-height ratio of P/e = 10. The experiments are carried out at the Reynolds numbers of 20,000, 40,000 and 80,000. The transient heat transfer test technique using thermochromic liquid crystals (TLCs) was applied to measure the detailed surface heat transfer distributions. The local mixed bulk temperatures (T mb ) are defined using the air thermocouples positioned in the centerline along the cooling passage. The static pressures are measured from the pressure taps machined as a part of the test model so that friction factors of the ribbed cooling passage are also evaluated. The friction factors are influenced by the characteristics of 3 dimensional converging duct of the experimental model. The full surface heat transfer coefficients, friction factors, and overall thermal performance of the cooling passage are reported for the comparison with the other cooling configurations.

The effects of heat transfer between the compressor structure and primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients, or so called “Bodie” transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Transient calculations with heat transfer show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation and the remainder is attributed to stage re-matching. It is found that heat transfer increases loading in the front stages and destabilizes the front block. Furthermore, the stage re-matching due to heat transfer alters the slope of the compressor characteristic and promotes modal-type stall inception. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects, such as NPSS, only capture 1.2 points of the 8.0 point stall margin reduction. This implies that, using this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Leakage flow between the rotating turbine blade tip and the fixed casing causes high heat loads and thermal stress on the tip and near the tip region. For this study, new squealer tips called partial cavity tips, which combine the advantages of plane and squealer tips, were suggested, and the effects of the cavity shape on the tip heat transfer coefficient and film cooling effectiveness were investigated experimentally in a low speed linear cascade. The suggested blade tips had a flat surface near the leading edge and a squealer cavity from the mid-chord to trailing edge region to achieve the advantages of both blade tip types. The heat transfer coefficient was measured via the 1-D transient heat transfer technique using an IR camera, and the film cooling effectiveness was obtained via the pressure sensitive paint (PSP) technique. Results showed that the heat transfer coefficient and film cooling effectiveness on the partial cavity tips strongly depended on the cavity shape. Near the leading edge, the heat transfer coefficients for the partial cavity tip cases were lower than that for the squealer tip case. However, the heat transfer coefficient on the cavity surface was higher for the partial cavity tip cases. The D10 tip showed a similar distribution of film cooling effectiveness to that of the PLN tip near the leading edge and the DSS tip near the mid-chord region. However, the overall averaged film cooling effectiveness of the DSS tip was higher than that of the D10 tip.

Prediction of the contact temperature between two materials in high speed rubbing contact is essential when analysing the wear mechanism and modelling wear rate for unlubricated contact. Conventionally the assumptions used for a pin-on-disc configuration are either a steady heat source at the contact for slow speed rotation or an annular source along the rubbing track at high rotating speeds. In this paper, the rotating heating source is solved using an in-house finite element method (FEM) code. This captures the full geometry and rotating speed of the rubbing bodies. The transient heat transfer problem is modelled in a quasi-static way: eliminating the computational cost of a transient 3D simulation. This reduced-order model is analytically shown to be suitable for contact temperature prediction over a wide range of rotating speeds, anisotropic thermal conductivity, and non-uniform thermal boundary conditions. The model calculates heat partition accurately for a thin rotating disc and short pin combination, which cannot be predicted using existing analytical solutions. The quasi-static numerical model and in-house FEM solver are validated against Ansys Mechanical and experimental measurements using infrared thermography. The numerical result demonstrates that the annular source assumption can significantly under-predict the contact temperature, especially at the rubbing interface. Explicit modelling of a thin disc results in a higher heat partition coefficient compared with the commonplace semi-infinite length assumption on both static and rotating components. Furthermore, the thermal anisotropy for bristle tufts is numerically evaluated, and the tuft-on-disc configuration is compared to the pin-on-disc configuration. Despite the effective thermal conductivity in the bristle tuft being approximately one order of magnitude lower than along the bristle length (treating the bristle as a porous medium), its impact on heat partition and contact temperature is shown to be limited.

Experimental investigation has been performed to study the film cooling characteristics of counter-inclined structures on the turbine vane leading edge. In this paper, four counter-inclined models are measured including cylindrical film holes with and without impingement holes, laid-back film holes with and without impingement holes. A semi-cylinder model is used to model the turbine vane leading edge. Two rows of film holes are located at ±15° on either side of the leading edge model, inclined 90° to the flow direction and 45° to the spanwise direction. Film cooling effectiveness and heat transfer coefficient have been obtained using a transient heat transfer measurement technique with double thermochromic liquid crystals with four blowing ratios ranging from 0.5 to 2 at a 1.0 density ratio. The results show that the film cooling effectiveness decreases with the increase of blowing ratio. No matter cylindrical hole or laid-back hole, the addition of impingement enhances the film cooling effectiveness. Compared with cylindrical hole, laid-back hole produces a better film cooling performance mainly because of stronger lateral momentum. Moreover, the benefits of both adding impingement and exit shaping are more obvious under a large blowing ratio.

Flight Environment Simulation Volume (FESV) is the most important part of Altitude Ground Test Facilities (AGTF). It’s temperature and pressure control precision determines the level of test ability of AGTF. Therefore, in order to study the temperature and pressure control problem of FESV and improve the modeling precision of FESV, the energy equation and gas state equation are used to deduce the temperature and pressure differential equations of FESV. Meanwhile, the heat transfer influence of FESV has been taken into account in this paper and the transient heat conduction of FESV is established by using a discretizing method. The temperature and pressure differential equations of FESV are linearized around a balance point and the uncertainty of actuators has been considered in multiplicative uncertainty. The augmented system of linear model of FESV and the actuators are obtained. For the sake of making the controller design and weighting function choice more easily, a normalization method is used to normalize the augmented system. For the purpose of achieving the temperature and pressure synchronic control of FESV, a two-degree-of-freedom integral type μ synthesis control design method is proposed. What’s more, for guaranteeing the designed μ synthesis controller has servo tracking and disturbance attenuation performance, the performance weighting functions are designed according to the frequency division weighting principle and the control weighting functions are designed by using the principle of low frequency free limit, medium frequency gradually increase the limit, and high frequency maximum limit. The MATLAB Robust Control Toolbox function dksyn is used to design the μ controller. In order to verify the effectiveness of designed μ controller, we assume two types of engine test conditions. The simulation results show, for the engine test condition one, the biggest relative tracking error of temperature is less than 0.5% and the relative steady state error of pressure is less than 0.1% and the relative tracking error of pressure slope signal is less than 3%. For the engine test condition two, the relative steady state error of temperature is less than 0.1% and the relative tracking error of temperature slope signal is less than 1%. To verify the advantage of designed μ controller, we designed a PID controller and compared the simulation results with μ controller. The comparison results showed that the designed μ controller provided better performance than the PID controller.

An experimental research of film cooling performance of three single dust-pan shaped hole rows in different positions of a turbine blade was carried out in the short-duration transonic linear cascade at stationary condition, which can model realistic engine aerodynamic conditions. The effects of inlet Reynolds number ( Re in = 2.5 × 10 5 ∼7.5 × 10 5 ), isentropic exit Mach number ( Ma is = 0.71∼0.91) and coolant blowing ratio ( M = 0.8∼2.6) on film cooling effectiveness are investigated. Three single hole rows are located at 11.7%, 36.3% and 55.6% relative arc on the pressure sides of three enlarged blade models respectively. The adiabatic film cooling effectiveness are derived from the surface temperatures based on transient heat transfer measurement method. The results show that in the range of blowing ratios studied in the present paper, for location 3 the cooling effectiveness decreases a lot with blowing ratio increasing due to the lift-off of coolant at high blowing ratios, while for location 1 and 2, the film cooling effectiveness increases with blowing ratio increasing, because the strong favorable pressure gradient and high concave curvature near the leading edge lead to a good attachment of coolant on the surface. At M ≤1.0 conditions, the film cooling effectiveness of location 1 and 2 is lower than that of location 3, which reflects that strong favorable pressure gradient and high concave curvature weaken film cooling performance at low blowing ratio conditions, while the effect is opposite when M is greater than 1.0. For location 1, the highest general cooling performance is obtained at Re in = 2.5 × 10 5 condition, and for location 2, the change of Re in has different effects on cooling effectiveness in different regions. In the range of Ma is studied in this paper, the change of Ma is has little effect on film cooling effectiveness.

This paper presents an experimental and numerical investigation on the film cooling with different coolant feeding channel structures. Two ribbed cross-flow channels with rib-orientation of 135° and 45° respectively and the plenum coolant channel have been studied and compared to find out the effect of rib orientation on the film cooling performances of cylindrical holes. The film cooling effectiveness and heat transfer coefficient were measured by the transient heat transfer measurement technique with narrow-band thermochromic liquid crystal. Numerical simulations with realizable k-ε turbulence model were also performed to analyze the flow mechanism. The results show that the coolant channel structure has a notable effect on the flow structure of film jet which is the most significant mechanism affecting the film cooling performance. Generally, film cooling cases fed with ribbed cross-flow channels have asymmetric counter-rotating vortex structures and related asymmetric temperature distributions, which make the film cooling effectiveness and the heat transfer coefficient distributions asymmetric to the hole centerline. The discharge coefficient of the 45° rib case is the lowest among the three cases under all the blowing ratios. And the plenum case has higher discharge coefficient than the 135° rib case under low blowing ratio. With the increase of blowing ratio, the discharge coefficient of the 135° rib case gets larger than the plenum case gradually, because the vortex in the upper half region of the coolant channel rotates in the same direction with the film hole inclination direction and makes the jet easy to flow into the film hole in the 135° rib case.

Heat transfer characteristics on the inner surface of a converging duct with ribs are investigated by conducting transient heat transfer experiments using thermochromic liquid crystals. In particular, features associated with the smooth duct are addressed as reference. The parameters studied are: the non-dimensional rib pitch ( p/w = 10, 20, 50, p is the rib pitch, w is the rib width), the rib blockage ratio ( e/H = 10%, 20%, e is the rib height, H is the channel height) and the inlet Reynolds number Re ( Re = 6000∼20000). The results are presented on the top wall and its center line. Due to the converging shape of channel, the value of Nu/Nu0 increases downstream the channel, especially near the side wall. The effects of these parameters are: Firstly, the heat transfer improves with the increasing rib height whose influence decreases with the rise of Re number. Secondly, the value of Nu/Nu0 rises firstly, reaching its maximum at p/w = 20, and then declines with an increasing rib pitch. Thirdly, the value of Nu/Nu0 rises firstly and decreases afterwards with the Re number, as the Re number is within the range of Re ≤12000.

The Transient Heat Transfer Facility (THTF) was developed at the University of Oxford to test full-scale high pressure compressor and turbine casing air systems using gas turbine engine representative secondary air system mass flow rates, total temperatures and total pressures. Transient casing response together with blade and disc responses governs achievable tip clearances in both compressors and turbines. Traditionally, casing thermal response is faster than disc response and so requires cooling. However, future engine designs may have thermal responses of the discs that are faster than the casings. In this paper we investigate the use of air impingement as a means to speed up the casing response to match such future designs. Two different impingement configurations were selected from a total of eight designs based on steady Reynolds-averaged Navier-Stokes (RANS) predictions of maximum heat transfer coefficient (HTC). The 3D thermal growth of the casing was characterised by the surface temperature rise over a given period of time to assess achievable dynamic response. The configurations were tested in the THTF as a rainbow set. The resulting casing metal temperatures were used in conjunction with an explicit solution of the conduction in the casing wall to calculate HTC due to impingement. The experimental set-up resembles a typical aircraft engine and therefore includes measurement uncertainties arising from features such as fixtures, seals, geometries and large surface areas that are then subjected to varying thermal inertias. These can lead to circumferential temperature non-uniformities, as evident from the experimental results. An uncertainty analysis was performed to quantify these effects on the overall casing thermal response. The experimental data was then compared against numerical predictions from an axisymmetric, 90° sector, conjugate heat transfer model of the facility using the two impingement plate designs. Overall, the values of HTC agreed well across the experimental and numerical results. Both approaches predicted differences in which of the two designs was more effective; however, advantages and limitations were identified in both approaches. The combined experimental and numerical study shows the significance of analysing the full annulus, at engine representative conditions and the benefit of an impingement array to potentially speed up casing response for future engines.

This paper experimentally investigates the film cooling performance of an enlarged turbine guide vane with full-coverage cylindrical hole film cooling in short duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and isentropic exit Mach number independently. The effects of mass flow rate ratio ( MFR =4.83%∼8.83%), inlet Reynolds number ( Re in = 1.7×10 5 ∼5.7×10 5 ), and isentropic exit Mach number ( Ma is =0.81∼1.01) are investigated. There are five rows of cylindrical film cooling holes on the pressure side and four such rows on the suction side respectively. Another four rows of cylindrical holes are provided on the leading edge to obtain a showerhead film cooling. The surface heat transfer coefficient and adiabatic film cooling effectiveness are derived from the surface temperatures measured by the thermocouples mounted in the middle span of the vane surface based on transient heat transfer measurement method. Mass flow rate ratio is shown to have a significant effect on film cooling effectiveness. The increase of mass flow rate ratio increases film cooling effectiveness on pressure side, while increasing this factor has opposite effect on film cooling effectiveness on the suction side. At the same mass flow rate ratio, increasing the Reynolds number can enhance the film cooling performance, the expectation is that at low mass flow rate ratio condition increasing the Reynolds number decreases film cooling effectiveness on the pressure side. The heat transfer coefficient increases with the mass flow rate ratio increasing on both pressure and suction side. At middle and high inlet Reynolds number condition, in the region of 0.4<s<0.6 on suction side, the coolant weakens heat transfer adversely.

A melting model which is based on the Enthalpy-Porous Medium method is applied to study the transient heat transfer and ice melting process in a multi-layered body which was covered by ice and heated by electrothermal de-icing system. Three typical cases including one-dimensional, two-dimensional and three-dimensional model are studied. The liquid fraction is computed based on the enthalpy balance. The mushy zone is modeled as a pseudo-porous medium in which the porosity ranges from 0 (solid) to 1 (liquid) as the ice melting. The melting time and process are both investigated as well as the ice-water interface profiles and water thickness for the three typical geometries. The computational results show that the melting model is valid for simulating the process of transient heat transfer and ice-melting. The Enthalpy-Porous Medium method can be used to study thermal de-icing systems of aircraft or aeroengine inlet. The method can also be used to judge the ice shedding. A parallel interface is introduced at the position of 0.2 mm away from the ice/abrasion shield interface to simulate ice shedding. The computational results show that the new melting model is effectiveness and can be used to provide detailed information for de-icing system design.

This study investigates on heat transfer enhancement in pin fin cooling channels. Experiments are conducted in a staggered pin fin array consisting of 15 rows. Heat transfer measurements are conducted in the pin fin cooling channel using the transient liquid crystal technique. The reference temperature is approximated by the fluid bulk temperature, acquired by thermocouples at specific positions. Thermal inertia of the used thermocouples is considered. One other problem that occurs while using relatively long thermocouples in short aspect ratio ducts is the heat conduction along the wires, the so called stem effect. This can lead to erroneous temperature measurements. The impact of the thermocouple immersion length on the temperature measurement is investigated. A detailed assessment of the space and time-wise varying temperature distribution is conducted for the appropriate reference temperature. This paper gives an overview about the experimental setup and the used transient measurement technique. Results are represented in terms of temperature distribution, heat transfer distribution and averaged Nusselt number at the endwall.

This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped pressure side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer results for the no coolant injection show a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction side as opposed to the pressure side in stationary studies can primarily be attributed to the effects of the blade relative motion.