Transpiration cooling represents the pinnacle of turbine cooling and is characterised by an intrinsic material porosity which achieves high internal convective cooling, and full coverage cooling films on the external surface subjected to the hot gases. Quasi-transpiration systems, such as the double-wall effusion system discussed here, attempt to replicate the cooling effect of transpiration systems. The double-wall system is characterised by a large internal wetted area providing high internal convective cooling performance, with a highly porous external wall allowing the formation of a protective film over the external surface. This paper presents a low-order thermal model of a double-wall system designed to rapidly ascertain cooling performance based solely on the geometry, solid thermal conductivity, and approximate surface heat transfer coefficients. The performance of the model is initially validated using experimental data with heat transfer coefficients for the low order model obtained from fully conjugate CFD simulations. Following this, a more controlled CFD study is undertaken with both fully conjugate and fluid only simulations performed on several double-wall geometries to ascertain both overall effectiveness and film effectiveness data. Data from these simulations are used as inputs to the low order thermal model developed and the results compared. The low order model successfully captures both the trends and absolute cooling effectiveness achieved by the various double-wall geometries. The model therefore provides an extremely powerful tool in which the cooling performance of double-wall geometries can be near instantaneously predicted during the initial design stage, potentially allowing geometry optimisation to rapidly occur prior to more in-depth, costly and time-consuming analyses of the systems being performed. This potential benefit is demonstrated via the implementation of the model with input boundary conditions obtained using empirical correlations.

Further improvements in gas turbine efficiency can be sought through more advanced cooling systems — such as the double-wall, effusion system — which provide high cooling effectiveness with low coolant utilisation. The double-wall system, as described here, comprises two walls, one with a regular array of impingement holes, the other with a closely-packed, regular array of film holes (characteristic of effusion systems). These walls are mechanically and thermally connected to one another via a bank of pedestals which increase coolant wetted area and turbulent flow features. However, a lack of data exists in the open literature on these systems. This study presents a novel experimental heat transfer facility designed with the intent of investigating flat plate versions of such double-wall geometries. Key features of the facility are presented including the use of recirculation to increase mainstream-to-coolant temperature ratio and the use of infrared thermography to obtain thermal measurements. Some rig commissioning characteristics are also provided which demonstrate well-conditioned, uniform flow. Both coolant and mainstream Reynolds numbers are matched to engine conditions, with Biot number within around 15% of engine conditions. The facility is used to assess the cooling performance of four double-wall effusion geometries which incorporate various geometrical features. Both overall effectiveness and film effectiveness measurements are presented at a range of coolant mass flows with conclusions drawn as to preferable features from a cooling perspective. The results from a fully conjugate CFD model of the facility are presented which utilised boundary conditions obtained during experimental runs. Additionally, a computationally efficient decoupled conjugate method developed previously by the authors was adapted to assess the experimental geometries with the results comparing favourably.

A detailed analysis of film cooling performance on a double-walled effusion-cooled blade is essential for both the coolant consumption optimization and assessment of the film to offer the desired levels of the turbine blade protection. Yet there are hardly any film effectiveness studies on double-wall full-coverage film cooled turbine blades. This paper presents a detailed film cooling effectiveness study over the full surface of a double-walled effusion-cooled high-pressure turbine rotor blade using Pressure Sensitive Paint (PSP). PSP permitted a non-intrusive and conduction-errors-free means of obtaining clean and distinct local distribution of film effectiveness on the blade surface making it possible to extract valuable film cooling effectiveness performance data on the whole blade surface. Three large-scale circular pedestal double-wall blade designs with varying pedestal height, pedestal diameter and cooling hole diameter were tested in a high-speed stationary single-blade linear cascade running at engine-representative Mach and Reynolds numbers. All the blades were tested within a range of representative modern engine coolant mass flow, ṁ c to mainstream, ṁ g ratios; 1.6% < ṁ c /ṁ ∞ < 5.5%. High porosity blade exhibited a better flow distribution and was found to consistently perform the best.

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.

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.

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.

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.

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.

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.

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.