A double-wall cooling scheme combined with effusion cooling offers a practical approximation to transpiration cooling which in turn presents the potential for very high cooling effectiveness. The use of the conventional conjugate CFD for the double-wall blade can be computationally expensive and this approach is therefore less than ideal in cases where only the preliminary results are required. This paper presents a computationally efficient numerical approach for analysing a double-wall effusion cooled gas turbine blade. An existing correlation from the literature was modified and used to represent the two-dimensional distribution of film cooling effectiveness. The internal heat transfer coefficient was calculated from a validated conjugate analysis of a wall element representing an element of the aerofoil wall and the conduction through the blade solved using a finite element code in ANSYS*. The numerical procedure developed has permitted a rapid evaluation of the critical parameters including; film cooling effectiveness, blade temperature distribution (and hence metal effectiveness) as well as coolant mass flow consumption. Good agreement was found between the results from this study and that from literature. This paper shows that a straightforward numerical approach that combines an existing correlation for film cooling from the literature with a conjugate analysis of a small wall element can be used to quickly predict the blade temperature distribution and other crucial blade performance parameters.

Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

The continuous drive for ever higher turbine entry temperatures is leading to considerable interest in high performance cooling systems which offer high cooling effectiveness with low coolant utilisation. The double-wall system discussed here, is an optimised amalgamation of more conventional cooling methods including impingement cooling, pedestals, and film cooling holes in a more closely packed array characteristic of effusion cooling. The system entails two walls, one with the impingement holes, and the other with the film holes. These are mechanically connected via the bank of pedestal thereby allowing conduction between the walls and increasing coolant wetted area and turbulent flow. However, in the open literature, data — and particularly experimental data — on such systems is sparse. This study presents a newly commissioned experimental heat transfer facility designed to investigate double-wall cooling geometries. The paper discusses some of the key features of the steady-state facility, including the use of infrared thermography to obtain overall cooling effectiveness measurements. The facility is designed to achieve both Reynolds and Biot (to within 10%) number similarity to those seen at engine conditions. The facility is used to obtain overall cooling effectiveness measurements for a circular pedestal, double-wall test piece at three coolant mass-flow conditions with the results presented and discussed. A fully conjugate CFD model of the facility was also developed providing greater insight into the internal flow field. Additionally, a computationally efficient, decoupled conjugate method developed by the authors for analysing such double-wall systems is run at conditions to match the experiments. The results of the simulations are encouraging, particularly given how computationally efficient the method is, with area-weighted, averaged overall effectiveness within a small margin of those obtained from the experimental facility.

When a gas turbine is shut down it cools asymmetrically due to natural convection. After some time, a thermal gradient develops across the rotor drum as the top section cools slower than the bottom. This induces a thermal stress that causes the shaft to deflect — a phenomenon known as rotor or thermal bow. In some cases, this can lead to compressor tip rubs. All engines are affected by rotor bow to some degree. However, the mechanisms of natural convective heat transfer and impact of design features on rotor bow is not well understood; thus, the issue is not identified until late in the design stage, hampering an engines development. A novel experimental technique and facility has been developed that allows for accurate, detailed measurements of natural convective heat transfer in a gas turbine annulus to be made. The apparatus has the capacity to test circumferential and axial variations in temperature distribution, in addition to an isothermal wall boundary. An IR camera looking directly at the rotor surface is used to provide high-resolution heat transfer measurements. The test piece for this study consisted of a cylinder section and a conical section. Results from the isothermal cylinder test cases agreed quite well with previously published data, which gives confidence to the method that has been employed. The addition of a conical section had a profound effect on the results; however, more work is required before a method can be developed which will accurately predict the heat transfer coefficients based on initial temperature distribution. A 2D concentric cylinder methodology has also been developed that gives results in strong agreement with previously published data. Going forward, this new facility and the CFD method will be further developed to help quantify the effect of various design features on the natural convective heat transfer performance of a large civil gas turbine.

The effect of film cooling on a high pressure turbine blade with an open squealer tip has been examined in a high speed linear cascade. The cascade operates at engine realistic Mach and Reynolds numbers, producing transonic flow conditions over the blade tip. Tests have been performed on two uncooled tip geometries with differing pressure side rim edge radii, and a cooled tip matching one of the uncooled cases. The pressure sensitive paint technique has been used to measure adiabatic film cooling effectiveness on the blade tip at a range of tip gaps and coolant mass flow rates. Complementary tip heat transfer coefficients (HTC) have been measured using transient infrared thermography, and the effects of the coolant film on the tip heat transfer and engine heat flux examined. The uncooled data show that the tip heat transfer coefficient distribution is governed by the nature of flow reattachments and impingements. The squealer tip can be broken down into three regions, each exhibiting a distinct response to a change in the tip gap, depending on the local behaviour of the overtip leakage flow. The edge radius of the pressure side rim causes the overtip leakage flow to change dramatically at low clearance. Complementary CFD shows that the addition of casing motion causes no further change on the pressure side rim. Injected coolant interacts with the overtip leakage flow, which can locally enhance the tip heat transfer coefficient compared to the uncooled tip. The film effectiveness is dependent on both the coolant mass flow rate and tip clearance. At increased coolant mass flow, areas of high film effectiveness on the pressure side rim coincide strongly with a net heat flux reduction and in the subsonic tip region with low heat transfer coefficient.

The drive to further increase gas turbine thermal efficiency and specific power output continue to elevate core temperatures well beyond the natural capabilities of the metals employed in their manufacture necessitating increasingly complex cooling systems. One such cooling mechanism is the double-wall, effusion-cooled system which combines in a very compact format, many cooling aspects already implemented in gas turbine cooling. To-date, thermomechanical stresses have provided one of the more significant challenges in the implementation of these systems and needs to be considered — alongside aerothermal performance — at the initial stages of design. This paper presents a novel computational method that has been developed to allow an integrated assessment of both the aerothermal and thermomechanical performance of double-wall cooling geometries. A decoupled conjugate method was developed in which internal cooling performance was ascertained via a conjugate CFD model in which the mainstream flow was not simulated. Instead, external film cooling performance was assessed via a superposition method that was developed and applied to a two-dimensionally varying correlation allowing streamwise film development to be modelled. Results of both the internal and external cooling simulations were then utilised in a conduction model to develop a complete thermal assessment of the geometry. The calculated temperature distribution was used in a thermomechanical FEA analysis permitting an insight into the stress field developed within the double-wall geometry under thermal load. The developed method was demonstrated in the assessment of seven circular pedestal, double-wall geometries in which a range of geometric parameters were investigated. The results provide an insight into the effect of varying these parameters on both the aerothermal performance of the selected geometries, along with the effect on the thermomechanical stress field developed.

Adiabatic film cooling effectiveness measurements are made on nozzle guide vane leading edges in an engine-realistic flow environment. The tested leading edges feature radial showerheads with different spanwise distributions of hole surface angle. The showerheads blow towards the midspan, except for one model with showerhead holes orthogonal to the vane surface. The results show that low surface angle radial showerhead holes generate high effectiveness within their rows and further downstream, but neglect the stagnation region lying between the two most upstream cooling hole rows. This downstream effectiveness gain is due to both the continued surface attachment of this coolant as it progresses downstream, and its beneficial interactions with downstream cooling jets. Moderate radial showerhead surface angles cause moderate coolant jet penetration into the mainstream, which promotes near-surface mixing of the coolant with the mainstream, increasing stagnation region effectiveness. The mixing effect is enhanced by the intense turbulence generated by combustor dilution jets. High surface angles may cause the stagnation region coolant to penetrate too far for either of these gains to be realised. Considering also the presence of endwall film cooling, these effects, taken together, suggest the superiority of radial showerheads which blow towards the midspan, as against those which blow towards each endwall. Surface temperature data is acquired by a novel infrared thermography technique which permits measurement of both heat transfer coefficient and film effectiveness from a single heated test.

This paper presents a computational investigation into the impact of cooling air injected through the stationary over-tip turbine casing on overall turbine efficiency. The high work axial flow turbine is representative of the high pressure turbine of a civil aviation turbofan engine. The effect of active modulation of the cooling air is assessed, as well as that of the injection locations. The influence of the through-casing coolant injection on the turbine blade over-tip leakage flow and the associated secondary flow features are examined. Transient (unsteady) sliding mesh simulations of a one turbine stage rotor-stator domain are performed using periodic boundary conditions. Cooling air configurations with a constant total pressure air supply, constant mass flow rate and actively controlled total pressure supply are assessed for a single geometric arrangement of cooling holes. The effects of both the mass flow rate of cooling air and the location of its injection relative to the turbine rotor blade are examined. The results show that all of the assessed cooling configurations provided a benefit to turbine row efficiency of between 0.2 and 0.4 percentage points. The passive and constant mass flow rate configurations reduced the over-tip leakage flow, but did so in an inefficient manner, with decreasing efficiency observed with increasing injection mass flow rate beyond 0.6% of the mainstream flow, despite the over-tip leakage mass flow rate continuing to reduce. By contrast, the active total pressure controlled injection provided a more efficient manner of controlling this leakage flow, as it permitted a redistribution of cooling air, allowing it to be applied in the regions close to the suction side of the blade tip which more directly reduced over-tip leakage flow rates and hence improved efficiency. Cooling air injected close to the pressure side of the rotor blade was less effective at controlling the leakage flow, and was associated with increased aerodynamic loss in the passage vortex.

This research focuses on film cooling of the trailing edge of a scaled up turbine rotor blade with engine-representative Mach number distribution. Pressure sensitive paint was used to obtain high-resolution adiabatic film cooling effectiveness measurements in the trailing edge region of the scaled turbine blade. The large scale, high-speed experimental set-up consists of a Perspex test section for maximum visibility of the PSP coated blade. The test section was designed to recreate a single blade passage of a gas turbine with inlet Mach and Reynolds numbers matching the corresponding values in an engine. The test blade has a constant cross section, representative of the mid-span profile of the high pressure turbine rotor blade. It was manufactured from aluminium to minimize temperature gradients over the surface of the test blade. In the current research, pressure surface cooling slots at the trailing edge were examined and the effect of cutback surface protuberance, or ‘land’, shapes on trailing edge film cooling was studied. Nitrogen and air were used as coolant gases giving a coolant to mainstream density ratio close to 1. Two land geometries-straight and tapered-were studied for a set of 6 blowing ratios from 0.4 to 1.4 in steps of 0.2. Land taper has a benefit for film cooling near the slot exit but its advantage reduces close to the trailing edge. For both geometries, film effectiveness falls with blowing ratio from 0.4 to 0.8 and increases with blowing ratio in the 0.8 to 1.4 range. Crossflow causes the coolant film to be biased towards one side of the lands. Film effectiveness results are compared with data from a scaled up low speed flat plat model of the trailing edge to explain the effect of acceleration on film cooling.

Extending disc life through online health monitoring has been a proven method of minimising engine downtime and maintenance costs. To properly monitor the disc requires a robust model of the disc’s non-linear thermal dynamics. A model can be improved by filtering the output using a measurement of the disc in real time. The damage models can then be computed with higher statistical confidence leading to increased safe life prediction. Recently, a model of disc temperature has been developed based on the proper orthogonal decomposition of simulated data. The model produced detailed thermal gradients for use in damage calculation and life assessment. This paper presents the development and implementation of a Kalman filter to augment that model. The location of the measurement has been assessed in order to select the most appropriate target for instrumentation. Points all around the front and back of the disc have been assessed, and the best practice result is found to be near the centre of the disc neck. Matching temperatures at this point represents a compromise between the fast dynamic response of the rim, with the slower response of the cob. The new model has been validated against an independent flight simulation that had previously been excluded from any training process. The addition of the Kalman filter allows the model to match aircraft dynamics outside the regular training trajectories. The accuracy is approximately ±30K, and there is a root-mean-square error of only 2K over the whole model at any one point in time.

This paper presents a novel approach to modelling disc temperature distributions for use in health monitoring applications. The method combines singular value decomposition techniques with a series of low order transfer functions to predict the thermal response over a typical flight cycle. The model inputs reflect typical online measurements; engine speed, altitude, and compressor exit temperature. The model’s complexity has been reduced by projection and truncation of the reference data and its decompositions. The model predicts transient temperatures anywhere within the disc’s profile to an accuracy of 20K, and similarly the thermal gradient can be calculated in any direction within the 2D cross-section. The result is a simple model that is suitable for application in a FADEC system, with the accuracy necessary to safely extend a disc’s life through online calculation of fatigue damage incurred for individual flights. The paper concludes by comparing the model with an earlier method using a grey box physics approach.

The continuous demand from the airlines for reduced jet engine fuel consumption results in increasingly challenging high pressure turbine nozzle guide vane (NGV) working conditions. The capability to reproduce realistic boundary conditions in a rig at the combustor-turbine interaction plane is a key feature when testing NGVs in an engine-representative environment. A large scale linear cascade rig to investigate NGV leading edge cooling systems has been designed with particular attention being paid to creating engine representative conditions at the inlet to the NGVs. The combustor simulator replicates the main features of a rich-burn design including large dilution jets and extensive endwall film cooling. A three-dimensional computational domain including the entire combustor simulator has been created and RANS CFD simulations have been run in order to match Reynolds number and mainstream-to-coolant momentum flux ratio; velocity and turbulence measurements have been acquired at the NGV inlet plane at ambient temperature. In this engine-representative environment the authors focused their attention on the flow field downstream of different endwall film cooling holes configurations: three arrangements of a double row of staggered cylindrical holes (lateral pitch-to-diameter ratio of 2–3–6) and one with intersecting holes (intersecting angle of 90°) are experimentally and numerically analyzed. Velocity, turbulence intensity and integral length scales are predicted and measured for a density ratio of 1 and coolant-to-mainstream momentum flux of 6. A hot wire sensor was mounted on a two-axis traverse mechanism able to move the probe in the spanwise and lateral directions. Three slots allowed to reposition the traverse and take measurements at three downstream locations (stream-wise distance-to-diameter ratio of 4.2–9.2–14.2). The research confirmed the strong influence of the endwall coolant on the flow field at the NGV inlet plane and the hole spacing results a key parameter in managing the film development. Closer-spaced hole configurations can assure an effective film coverage. The integral length scales are strongly connected to the hole diameter and spacing. Intersecting holes can potentially reduce the amount of required coolant at a fixed pressure ratio, but they offer worst film performance than cylindrical holes. RANS simulations proved to be able to get the main trends shown by the measurements.

A transient method of measuring the local heat transfer under an array of impinging jets has been developed. The use of a temperature sensitive coating consisting of three encapsulated thermochromic liquid crystal materials has allowed the calculation of both the local adiabatic wall temperature and the local heat transfer coefficient over the complete surface of the target plate. The influence of the temperature of the plate through which the impingement gas flows on the target plate heat transfer has been quantified. Results are presented for a single inline array configuration over a range of jet Reynolds numbers.

The heat transfer coefficient over the surface of a pedestal with fillet radii has been measured using thermochromic liquid crystals and the transient heat transfer method. The tests were performed at engine representative Reynolds numbers for a geometry typical of those used in turbine blade cooling systems. The heat conduction process that occurs in the engine was subsequently modelled numerically with a finite element discretization of the solid pedestal. The measured heat transfer coefficients were used to derive the exact boundary conditions applicable to the engine. The temperature field within the pedestal, calculated using the correct heat transfer coefficient distribution is compared to that calculated using an area averaged heat transfer coefficient. Metal temperature differences of 90 K are predicted across the blade wall.

A new method of processing the liquid crystal colour change data obtained from transient heat transfer experiments is presented. The approach uses the full intensity history recorded during an experiment to obtain an accurate measurement of the surface heat transfer coefficient at selected pixels. Results are presented for a model of a turbine blade cooling passage with combined ribs and film cooling holes. The implementation of the technique and the advantages to be gained from its application are discussed.