Prevention of high cycle fatigue failures of blades caused by increased level of vibrational stresses is important problem in engine design and development. Different types of dampers are usually installed under the blade platforms in the region between the shanks of adjacent blades are usually used to reduce the vibrational stresses in the turbine blades. The action of these dampers is based on the vibrational energy dissipation by the dry friction forces arising due to the relative displacement of the contacting surfaces of the damper and the blade. Adoption of a reliable choice of damper design mass and stiffness at the design stage of the aircraft engines requires an experimentally validated estimation of its damping efficiency and its influence on the blade frequency characteristics. The dampers need to be tested under conditions that are maximally close to the operational conditions. This paper presents an analysis of the stress gauging results of a turbine wheel with box-type dampers of various mass. The tests were carried out on a dynamic bench where resonance oscillations of a stage for a given mode were excited by one of the harmonics to a given rotational speed under the corresponding forces pressing the dampers onto the blade platform. An estimation of the damper efficiency in a wider range of pressing forces was also obtained by conducting the laboratory tests using on a specially designed vibration unit that simulates the dampers effect on a set of three blades. Numerical analysis of the damper mass and stiffness influence on the frequency and damping characteristics of a real turbine wheel was performed. Obtained experimental data and the GAP model of contact interaction of the insert with blades were used in the MSC. Nastran software package. Analysis showed a satisfactory correlation between the calculated and experimental estimates of the decrease in resonant stresses and resonance frequency changes when the dampers are inserted. The results presented here confirm the possibility of increasing the reliability of the choice of mass and stiffness of turbine blade dampers at the design stage.

In the context of an engine design calculation, isentropic or polytropic efficiencies of turbomachinery components are assumed at the outset of the cycle analysis and their values are updated or validated following the aerodynamic design of the components. In the present paper, aerodynamic design calculations of axial-flow compressors and turbines are directly integrated into the corresponding performance component models. This creates a consistent, single-step preliminary design and performance modelling process using a relatively small number of physical and geometric inputs. The aerodynamic design for establishing a component’s overall efficiency is accomplished through a mean-line, stage-by-stage approach where the stagewise isentropic efficiency is calculated employing either loss or semi-empirical correlations. From this process, the stagewise flow annulus radii are also obtained and are used to axially size the component stages assuming the blade aspect ratio and axial gapping distributions. The component flowpath geometry is then produced by simply “stacking” axially the component stages. The developed method is validated against publicly available data for a high-pressure compressor and a low-pressure turbine. Finally, the effectiveness of the method is demonstrated by considering the multi-point design of a High Bypass Ratio Geared Turbofan Engine with bypass Variable Area Nozzle.

This work details an analytical assessment of heat generation in a turbine aero-engine main-shaft bearing and the development of a model to predict that heat generation. The new model is based on an empirical model, previously developed by the Air Force Research Laboratory (AFRL), which features physics based terms multiplied by empirical regression coefficients. That model proved to be limited in that portions of the terms were essentially an extension of the regression coefficients due to the fact that the experimental data was limited to that of one bearing. Additionally, there were separate models for each rolling element material. To develop the new model, the validated bearing analysis code ADORE was used to generate power loss data for angular contact ball bearings of various sizes. The effects of speed, thrust load, pitch diameter, element diameter, number of rolling elements, lubricant inlet temperature, lubricant flow rate, and rolling element material (AISI M50 bearing steel and silicon nitride) are examined. Speed and thrust load are addressed at four levels each. Number of elements, bore diameter, and element diameter as well as lubricant temperature and flow rate are each addressed at three levels. These effects are captured in the model through traction (friction), churning (drag), and shearing (viscous) terms and their respective regression coefficients. The material effect is address through the use of an effective elastic modulus within an estimate of raceway to rolling element contact area. The performance of the model was then compared with experimental data collected in the AFRL High Mach Engine (HME) Bearing Rig. The model created in this work provides designers with an effective tool to examine bearing heat generation during the early engine design phases, avoiding the significant computational and front end expense of other, more detailed methods.

The effects of high shaft power offtake in a direct drive, a geared drive, and a novel turbofan configuration are investigated. A design and off-design performance analysis shows the configuration specific limitations and advantages. The more electric aircraft (MEA) concept promises to offer advantages with respect to aircraft performance, maintenance and operating costs. The engines for the MEA concept are based on conventional turbofan architectures. These engines are designed for significantly increased shaft power offtake that is required by the airframe, and the shaft power is usually taken off the high-pressure spool. This can impair the off-design performance of the engine and lead to compromises during engine design and to operability limitations. Taking the power off the low-pressure spool mitigates some of the problems but has other limitations. In this work, an alternative novel turbofan architecture is investigated for its potential to avoid the problems related to high shaft power offtakes. This architecture is called the dual drive booster because it uses a summation gearbox to drive the booster from both the low- and high-pressure spool. The shaft power, if taken off the booster spool, is effectively provided by both the low- and high-pressure spools, which allows the provision of very high power levels. This new concept is benchmarked against a two-spool direct drive and a geared drive turbofan. Furthermore, it is described, how the new architecture can incorporate an embedded motor generator. The presented concept mitigates some of the problems which are encountered during high power offtake in conventional configurations. In particular, the core compressors are less affected by a change in shaft power offtake. This allows higher power offtakes and gives more flexibility during engine design and operation. Additionally, the potential to use the new configuration as a gas turbine-electric hybrid engine is assessed, where electrical power boost is applied during critical flight phases. The ability to convert additional shaft power is compared with conventional configurations. Here, the new configuration also shows superior behavior because the core compressors are significantly less affected by power input than in conventional configurations. The spool speed and its variation is more suitable for electrical machines than in conventional configuration with low-pressure spool power transfer. The dual drive booster concept is particularly suited for applications with high shaft power offtakes and inputs, and should be considered for propulsion of more electric aircrafts.

Aerobatic aircraft have become popular for the training of military pilots, and nowadays an increasing number of such airframes are being developed. Modern turboprop engines provide high performance allowing the pilots to get similar handling characteristics to military jet aircraft engines. Prior to the availability of high performance turboprops, the basic pilot training was conducted using jet aircraft. Furthermore, the introduction of electronic control systems on last-generation turboprop engines enables single lever control, making it an ideal candidate for this type of aerobatic and training airframes. This new type of engine operation is however accompanied by several challenges from the point of view of the engine design and installation aerodynamics. GEAC has gone through a complex design process, in cooperation with the airframer, to validate the design of a new aerobatic aircraft inlet in the context of developing an aerobatic version of the H80 engine. In order to ensure a) surge-free operation, b) optimal engine performance and c) effective ice/FOD separation in inclement weather conditions and during any kind of aerobatic maneuver, the team has done extensive CFD predictions of the flow behaviour, performance/operability studies and finally a ground test campaign. First, a back-to-back comparison of the aerobatic inlet geometry versus a reference commuter inlet geometry was conducted. Then, flight conditions were simulated in calm and crosswind environments. Distortion patterns were examined using in-house developed tools and the diverse sources of distortion were identified. One of the results is the introduction of geometry improvements to guarantee improved performance and extended engine operability range. Advanced propeller modeling techniques were introduced and benchmarked in order to have the most exact representation of the propeller aerodynamic effect on inlet flow. Finally, a test campaign was conducted for validation purposes. An exhaustive instrumentation, data acquisition system and detailed test program were developed to validate CFD methods and assumptions made during the design phase, and to raise our confidence in the flight conditions simulation results.

The numerical description of combustion provides a good understanding of the thermal boundary conditions of a combustor already in the early design phase. With an accurate simulation, the thermal load on the combustor wall as well as the turbine inlet profile can be estimated and potentially tailored to the structural requirements. At MTU Aero Engines AG, the authors set up a multi-physics solver in OpenFOAM to support the development process for combustors. This article shows the activities for a redesign of a commercial engine combustor and the successful numerical description of all related physical aspects. The comparison of combustor liner temperatures predicted via Conjugate Heat Transfer (CHT) with thermal paint results show excellent agreement. The numerical results allow the evaluation of different redesign options with respect to maximum liner temperatures and impact on turbine inlet temperature profiles. The application of a multi-physics solver to real engine combustor designs is limited to the availability of computational power and acceptable running time. The applicability of the presented approach to complex engine designs is not only shown by the successful simulation results. The choice of an open source code allows the usage of massive parallelization. A parallelization up to 1024 CPUs was tested to evaluate the potential of speeding up the simulations. The results are encouraging and show that complex meshes with ∼15 million cells can be simulated within acceptable calculation time.

At the preliminary design stage of the engine design process, the behaviour and efficiency of different engine designs are investigated and evaluated in order to find a best matching design for a set of engine objectives and requirements. The prediction of critical part temperatures as well as the reduction of the uncertainty of these predictions is decisive to bid a competitive technology in aerospace technology. Automated workflows and Design of Experiments (DOE) are widely used to investigate large number of designs and to find an optimized solution. Nowadays, technological progress in computational power as well as new strategies for data handling and management enables the implementation of large DOEs and multi-objective optimizations in less time, which also allows the consideration of more detailed investigations in early design stages. This paper describes an approach for a preliminary-design workflow that implements adaptive modelling and evaluation methods for cavities in the secondary air system (SAS). The starting point for the workflow is a parametric geometry model defining the rotating and static components. The flow network within the SAS is automatically recognized and CFD and Thermal-FE models are automatically generated using a library of generic models. Adaptive evaluation algorithms are developed and used to predict values for structural, air system and thermal behaviour. Furthermore, these models and evaluation techniques can be implemented in a DOE to investigate the impact of design parameters on the predicted values. The findings from the automated studies can be used to enhance the boundary conditions of actual design models in later design stages. A design investigation on a rotor-stator cavity with axial through flow has been undertaken using the proposed workflow to extract windage, flow field and heat transfer information from adiabatic CFD calculations for use in thermal modelling. A DOE has been set up to conduct a sensitivity analysis of the flow field properties and to identify the impact of the design parameters. Additionally, impacts on the distribution of the flow field parameters along the rotating surface are recognized, which offers a better prediction for local effects in the thermal FE model.

The effects of upstream leakage injection angle on film cooling effectiveness of a turbine vane endwall with various endwall film-hole designs were examined by applying PSP measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create additional controllable upstream leakage flow to cool the vane endwall. Three potential leakage injection angles were studied: 30°, 40°, and 50°. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, mid-chord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream leakage injection angle.

The optimization and evaluation of blading clearance is important for gas turbine efficiency and performance. The Ansaldo GT36 gas turbine offers high efficiency together with outstanding flexibility across a large load range. Active management of engine clearances during the complete development process followed by a thorough validation on the Ansaldo test plant facility in Birr, Switzerland enables the GT to attain ambitious clearance targets. The clearance at baseload must be minimized but is limited by the pinch point clearance during cold, warm and hot start-ups — including normal and fast ramp-up and/or shutdown. Therefore transient analysis is necessary for covering the different operating conditions. A well-established process of 2d finite element modelling of the whole engine model (WEM) comprised of axis-symmetric and plane stress elements was used during the design process from concept to detailed design to optimize the clearances. It delivers the transient stator and rotor deformation and together with the compressor and turbine airfoil deformation based on 3D models the basic clearance evaluation process is defined. The GT engine design was significantly influenced, starting with a simplified version of the WEM for identification of the most promising variants. Subsequently a detailed WEM was developed which is fully validated against measurements on the test engine. Different 3D effects are considered separately at identified critical transient conditions and overlaid on the 2d clearances which lead to the final optimized clearances. In addition to this, limitations from each step of the manufacturing process were identified and improved to reduce tolerances and uncertainties to their minimum. The results of the calculation and clearance prediction process are compared against clearance measurements during all kinds of GT operation and cooldown. Passive clearance indicators showing the remaining gap till rubbing would occur and rub marks, in areas that tolerate it, further validate the clearances and clearance prediction process.

This paper presents a practical approach to designing a gas turbine nozzle with the help of the Aircraft Engine Design textbook as well as the software program Nozzle, a subprogram within the Aircraft Engine Design System Analysis Software suite AEDsys. The current textbook and software allow for a variable wetted length of the converging and diverging nozzle sections. Critical feedback from industry experts has inspired an attempt to design a nozzle with fixed wetted material lengths. This paper is written to augment classroom treatment, but will also support others who use the Aircraft Engine Design text and software for a preliminary engine design capstone. This approach is further guided by the actual scaling of the Pratt & Whitney F100 variable geometry converging-diverging nozzle, where wetted lengths are fixed. The chief goal is to equip students at the United States Air Force Academy with a refined approach that is more realistic of a manufactured nozzle design, producing a graphical representation of a nozzle schedule at different speed and altitude flight conditions, both with and without afterburner.

This paper describes a new engine-parts facility at the University of Oxford for high technology-readiness-level research, new technology demonstration, and for engine component validation. The Engine Component AeroThermal (ECAT) facility has a modular working section which houses a full annulus of engine components. The facility is currently operated with high-pressure nozzle guide vanes from a large civil jet-engine. A high degree of engine similarity is achieved, with matched conditions of Mach number, Reynolds number, and coolant-to-mainstream pressure ratio. For combustor-turbine interaction studies, a combustor simulator module is used, which is capable of both rich-burn and lean-burn combined temperature, swirl and turbulence profiles. The facility is being used for aerothermal optimisation research (e.g., novel cooling systems, aerodynamic optimisation problems, capacity sensitivity studies), computational fluid dynamics validation (aerodynamic predictions, conjugate predictions), and for component validation to accelerate the engine design process. The three key measurement capabilities are: capacity characteristic evaluation to a precision of 0.02%; overall cooling (metal) effectiveness measurements (using a rainbow set of parts if required); and aerodynamic loss evaluation (with realistic cooling, trailing-edge flow etc.). Each of these three capabilities have been separately developed and optimised in other facilities at the University of Oxford in the last 10 years, to refine aspects of facility design, instrumentation design, experimental technique, and theoretical aspects of scaling and reduction of experimental data. The ECAT facility brings together these three research strands with a modular test vehicle for rapid high technology-readiness-level research, demonstration of new technologies, and for engine component validation. The purpose of this paper is to collect in one place — and put in context — the work that led to the development of the ECAT facility, to describe the facility, and to illustrate the accuracy and utility of the techniques by presenting typical data for each of the key measurements. The ECAT facility is a response to the changing requirements of experimental turbomachinery testing, and it is hoped this paper will be of interest to engine designers, researchers, and those involved in major facility developments in both research institutes and engine companies.

The flow through radial turbines in automotive turbocharging applications is inherently highly unsteady, exhibiting high amplitude pulsations of varying frequencies. This can have a significant effect on the performance of the turbine, and therefore accurate modelling of these effects is key in good engine design, particularly in the realm of 1-D modelling. Assuming that the turbine stage acts in a quasi-steady manner can result in modelling errors. To better capture the physical effects of unsteady flow, a novel 1-D model of a radial turbine volute was developed, within the engine modelling software GT-Power. A meanline model was developed and integrated into the volute model as a means of modelling the rotor, which was assumed to be quasi steady. Experimental work was done to characterise the unsteady performance of the turbine, and the results were used to validate the model. The model showed good agreement with experimental data as well as some small improvements compared to industry standard methods of modelling.

It has been recognised in recent years that high altitude atmospheric ice crystals pose a threat to aircraft engines. Instances of damage, surge and shutdown have been recorded at altitudes significantly greater than those associated with supercooled water icing. It is believed that solid ice particles can accrete inside the core compressor, although the exact mechanism by which this occurs remains poorly understood. Development of analytical and empirical models of the ice crystal icing phenomenon is necessary for both future engine design and this-generation engine certification. A comprehensive model will require the integration of a number of aerodynamic, thermodynamic and mechanical components. This paper studies one such component, specifically the thermodynamic and mechanical processes experienced by ice particles impinging on a warm surface. Results are presented from an experimental campaign using a heated and instrumented flat plate. The plate was installed in the Altitude Icing Wind Tunnel (AIWT) at the National Research Council of Canada (NRC). This facility is capable of replicating ice crystal conditions at altitudes up to 9 km and Mach numbers up to 0.55 [1]. The heated plate is designed to measure the heat flux from a surface at temperatures representative of the early core compressor, under varying convective and icing heat loads. Heat transfer enhancement was observed to rise approximately linearly with both total water content and particle diameter over the ranges tested. A Stokes number greater than unity proved to be a useful parameter in determining whether heat transfer enhancement would occur. A particle energy parameter was used to estimate the likelihood of fragmentation. Results showed that when particles were both ballistic and likely to fragment, heat transfer enhancement was independent of both Mach and Reynolds numbers over the ranges tested.

The purpose of this paper is to explore GasTurb 12, a commercial gas turbine engine performance simulation program, for supplementary use on an introductory propulsion design project in an undergraduate course. This paper will describe several possible opportunities for supplementing AEDsys (Aircraft Engine Design System Analysis) version 4.012, the engine design software tool currently in use. The project is assigned to juniors taking their first propulsion course in the aeronautical engineering major at the USAF Academy. This course, Aeronautical Engineering 361, which focuses on cycle analysis and selection, is required of all aero majors and is used to satisfy the ABET Program Criterion requiring knowledge of propulsion fundamentals. This paper describes the most recent design project that required the students to re-engine the USAF T-38 with the aim of competing for the Advanced Pilot Training Program (T-X) program. The goal of the T-X program is to replace the T-38 aircraft that entered service in 1961 with an aircraft capable of sustained high-G operations that is also more fuel efficient. The design project required the students to select an engine-cycle for a single, non-afterburning, mixed stream, low bypass turbofan engine to replace the two J85 turbojets currently in the T-38. It was anticipated that the high specific thrust requirements might possibly be met through the use of modern component measures of merit to include a much higher turbine inlet temperature. Additionally, it was anticipated that the required 10% reduction in thrust specific fuel consumption might possibly be achieved by using a turbofan engine cycle with a higher overall pressure ratio. This paper will describe the use of GasTurb 12 to perform the same design analysis that was described above using AEDsys as well as additional features such as numerical optimization, temperature-entropy diagrams, and the generation of scaled, two-dimensional engine geometry drawings. The paper will illustrate how GasTurb 12 offers important supplementary information that will deepen student understanding of engine cycle design and analysis.

Due to a high degree of complexity and computational effort, overall system simulations of jet engines are typically performed as 0-dimensional thermodynamic performance analysis. Within these simulations and especially in the early cycle design phase, the usage of generic component characteristics is common practice. Of course these characteristics often cannot account for true engine component geometries and operating characteristics which may cause serious deviations between simulated and actual component and overall system performance. This leads to the approach of multi-fidelity simulation, often referred to as zooming, where single components of the thermodynamic cycle model are replaced by higher-order procedures. Hereby the consideration of actual component geometries and performance in an overall system context is enabled and global optimization goals may be considered in the engine design process. The purpose of this study is to present a fully automated approach for the integration of a 3D-CFD component simulation into a thermodynamic overall system simulation. As a use case, a 0D-performance model of the IAE-V2527 engine is combined with a CFD model of the appropriate fan component. The methodology is based on the DLR in-house performance synthesis and preliminary design environment GTlab combined with the DLR in-house CFD solver TRACE. Both, the performance calculation as well as the CFD simulation are part of a fully automated process chain within the GTlab environment. The exchange of boundary conditions between the different fidelity levels is accomplished by operating both simulation procedures on a central data model which is one of the essential parts of GTlab. Furthermore iteration management, progress monitoring as well as error handling are part of the GTlab process control environment. Based on the CFD results comprising fan efficiency, pressure ratio and mass flow, a map scaling methodology as it is commonly used for engine condition monitoring purposes is applied within the performance simulation. Hereby the operating behavior of the CFD fan model can be easily transferred into the overall system simulation which consequently leads to a divergent operating characteristic of the fan module. For this reason, all other engine components will see a shift in their operating conditions even in case of otherwise constant boundary conditions. The described simulation procedure is carried out for characteristic operating conditions of the engine.

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.

Multi-row film cooling is widely used on both suction side and pressure side of turbine vane, and the coolant behavior is considerable for engine design. Main work of this paper is to find out the accuracy of superposition predictions. Experiments were conducted on flat plates with double rows of cooling holes. The method of stable infrared measurement technique was used to measure surface temperature. Four factors, including hole shape, hole arrangement, row-to-row spacing and blowing ratio were simulated. Numerical simulation using commercial software ANSYS Fluent was also performed to observe the flow structure and film cooling mechanisms between each row. Result showed that the blowing ratio within the range of 0.5 to 2 has an obvious influence on the accuracy of superposition prediction. At low blowing ratio, results obtained by superposition method agreed well with the experimental data while the increase of blowing ratio caused a decrease in accuracy. Another significant factor is hole arrangement, results obtained by superposition prediction was nearly the same as experimental values on staggered arrangement plates while it was much higher on in-line arrangement plates. For different hole shapes, the accuracy of superposition prediction on converging-expanding holes was better than cylinder holes and compound angle holes. For both two hole spacing in this paper, prediction results show good agreement with the experiment results.

This study focuses on specifics of gas turbine film cooling. Laboratory film cooling tests are important for industry because actual engine conditions are too hot, too small, and too fast to take accurate and high resolution measurements. Experiments are typically conducted using a plenum to feed coolant through round or shaped film cooling holes. Less common are experiments using cross-flow fed coolant, a method that flows coolant perpendicular to the mainstream flow and better represents engine designs. There are a few studies that have explored shaped holes in cross-flow, but none have looked at the effect cross-flow channel parameters other than Mach number. Here, the effectiveness of film cooling is quantified by measuring adiabatic effectiveness on a flat plate with a single row of shaped film cooling holes in cross-flow. A preliminary examination of the effect of cross-flow versus plenum fed coolant on the adiabatic effectiveness of the axial 7-7-7 shaped hole, a laidback fan-shaped hole with a 30 degree injection angle, is first conducted. Subsequently, the effects of two internal coolant parameters on film cooling effectiveness are presented: Reynold’s number inside the cross-flow channel, and velocity ratio (defined as the ratio of cross-flow channel average velocity to mainstream velocity). By measuring the effect of these parameters, a chain of relative importance can be generated and applied to future experimentation. Parameters that heavily influence film cooling effectiveness can be studied further and optimized for turbine film cooling design.

The continuing maturation of metal laser-sintering technology (DMLS) presents the opportunity to de-risk the engine design process by experimentally down-selecting HPNGV cooling designs using laboratory tests of laser-sintered — instead of cast — parts to assess thermal performance. Such tests could be seen as supplementary to thermal-paint-test engines, which are used during certification to validate cooling system designs. In this paper, we compare conventionally cast and laser-sintered titanium alloy parts in back-to-back experimental tests at engine-representative conditions over a range of coolant mass flow rates. Tests were performed in the University of Oxford Annular Sector Heat Transfer Facility. The thermal performance of the cast and laser-sintered parts — measured using new infra-red processing techniques — is shown to be very similar, demonstrating the utility of laser-sintered parts for preliminary engine thermal assessments. We conclude that the methods reported in this paper are sufficiently mature to make assessments which could influence engine development programmes.

This paper systematically presents a complete leakage comparison for various types of wear experienced by labyrinth seals. Labyrinth seals used in turbine engines are designed to work at a clearance during steady-state engine operations. The tooth tip rubs the stator and wears either itself or the stator surface during transient operations, depending on the material properties of the tooth and stator. Any type of wear that increases clearance or deforms the tooth tip will cause permanent and unpredictable leakage degradation. This negatively affects the engine’s overall efficiency, durability, and life. The teeth have been reported to wear into a mushroom profile or into a rounded profile. A rub-groove on the opposing surface may form in several shapes. Based on a literature survey, five rub-groove shapes are considered in this work. They are: rectangle, trapezoid (isosceles and acute), triangle, and ellipse. In this work, leakage degradation due to wear is numerically quantified for both mushroomed and rounded tooth wear profiles. It also includes analyses on rounded teeth with the formation of five rub-groove shapes. All parameters are analyzed at various operating conditions (clearance, pressure ratio, number of teeth, rotor speed). CFD analyses are carried out by employing compressible turbulent flow in a 2-D axi-symmetrical coordinate system. CFD analyses show that the following tooth-wear conditions affect leakage from least to greatest: unworn, rounded, and mushroomed. These are for an unworn flat stator. It is also observed that rub-groove shapes considerably affect the leakage depending on the clearance. Leakage increases with the following groove profiles: triangular, rectangular, acute trapezoidal, isosceles trapezoidal, and elliptical. The results show that any type of labyrinth seal wear has significant effects on leakage. Therefore, leakage degradation due to wear should be considered during the engine design phase.